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Shahil Kanji
Master of Applied Science
Graduate Department of Aerospace Science and Engineering
University of Toronto
2015
NORSAT-1 is a multi-payload microsatellite mission funded by the Norwegian Space Center,
with three overall objectives: investigating solar radiation, space plasma research, and
developing improved methods for detection and management of ship traffic. The successful
development of the NORSAT-1 platform aims to lay the groundwork for additional low-cost
microsatellites in the NORSAT series, and expand the Norwegian presence in space and space-
based ship tracking technologies.
This thesis provides some insight into the NORSAT-1 satellite platform design, and focuses
heavily on the mechanical aspects of design, analysis, and testing. The structural design is
detailed from the early conceptual design phases, and follows the development to the
manufacturing, integration, and testing of the flight spacecraft. Validation of the design through
finite element modeling is presented, along with the development and design of two honeycomb
composite solar panels, and two deployable whip antennas.
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First thanks goes out to my supervisor, Dr. Robert E. Zee, for giving me the opportunity to
pursue a MASc. degree at the Space Flight Lab (SFL). Not really knowing what I was getting
into, my position at SFL turned out to be exactly what I was looking for. It has truly been a
rewarding and challenging experience working here, with amazingly knowledgeable people and
on amazingly awesome technology.
Special thanks to my manager Alex Beattie, who never ceased to entrust me with large project
responsibilities, allowing me to truly learn, grow, and develop my engineering skills, and get the
most out of my masters degree. Thanks to all of the staff and students at SFL for providing such
a welcoming and supportive environment, and especially to the many mechanical apt persons,
Cordell Grant, Benoit Larouche, Dumitru Diaconu, Ben Risi, Mike Ligori, and Brent Brakeboer
for providing much guidance and wisdom to many of the engineering problems with which I was
faced. Thanks to the rest of my fellow students at SFL and UTIAS as well for rendering these
past two years manageable, in particular the students of my year.
A small thanks goes out to my roommate at the time of writing this thesis who graciously offered
to proofread this 100 page monster, and then only got halfway. Thanks also to all of my close
friends for dealing with my constant bailing of weekend plans over the summer while I was
writing this. Final thanks goes out to my family, in particular my parents, who have supported
me my entire life, giving me the freedom and encouragement to pursue what makes me happy.
And thanks to my siblings for always pushing me to be better than the curve, and become
something great.
As the saying goes, enjoy the following.
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Acknowledgments ........................................................................................................................ iii Table of Contents ......................................................................................................................... iv List of Tables ................................................................................................................................ vi List of Figures .............................................................................................................................. vii
Introduction .............................................................................................................................. 1 11.1 Space Flight Laboratory ...................................................................................................... 1 1.2 The NORSAT-1 Microsatellite ........................................................................................... 2
1.3 Thesis Objectives and Outline ............................................................................................ 3
The NORSAT-1 Microsatellite................................................................................................ 5 22.1 Mission Overview ............................................................................................................... 5
2.2 Capabilities ......................................................................................................................... 5
2.2.1 Structure .................................................................................................................. 6
2.2.2 Telemetry and Command ........................................................................................ 6 2.2.3 Thermal ................................................................................................................... 6
2.2.4 Attitude and Control ............................................................................................... 7 2.2.5 Command and Data Handling ................................................................................. 7
2.2.6 Power ...................................................................................................................... 7
2.3 Payloads .............................................................................................................................. 8
2.3.1 Compact Lightweight Absolute Radiometer .......................................................... 9
2.3.2 Langmuir Probes ................................................................................................... 11
2.3.3 AIS Receiver ......................................................................................................... 13
Structural Subsystem Design ................................................................................................ 16 33.1 Driving Requirements ....................................................................................................... 16
3.2 Design Concept ................................................................................................................. 19
3.2.1 Initial Project Status .............................................................................................. 19 3.2.2 Current Design and Layout ................................................................................... 20
3.2.3 Electromagnetic Interference Mitigation .............................................................. 27 3.3 Design for Thermal ........................................................................................................... 29
3.3.1 Solar Panel Wings ................................................................................................. 30 3.3.2 Solar Cell Isolation ............................................................................................... 30
3.3.3 Thermal Surface Additions ................................................................................... 31
3.4 Payload Accommodations ................................................................................................ 33 3.4.1 CLARA ................................................................................................................. 33
3.4.2 AIS Receiver ......................................................................................................... 34
3.4.3 Langmuir Probes ................................................................................................... 35
3.5 Design for Assembly and Disassembly ............................................................................ 35 3.6 Wiring Harness Development ........................................................................................... 37
3.6.1 Solid Model Wiring .............................................................................................. 39
3.7 Materials Selection ............................................................................................................ 41 3.8 Mass Budget ...................................................................................................................... 42
3.9 Solid Modeling .................................................................................................................. 42 3.10 Design Evolution ............................................................................................................. 43
3.11 Manufacturing ................................................................................................................. 44
Finite Element Analysis ......................................................................................................... 45 44.1 Modeling ........................................................................................................................... 45
4.1.1 Primary Structure .................................................................................................. 45
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4.1.2 Solar Panel Wings ................................................................................................. 46 4.1.3 Components and Connections ............................................................................... 47
4.2 Boundary conditions ......................................................................................................... 48
4.2.1 Constraints ............................................................................................................ 48
4.2.2 Loads ..................................................................................................................... 48 4.3 Results ............................................................................................................................... 49
4.3.1 Stress Analysis ...................................................................................................... 49
4.3.2 Displacement Analysis .......................................................................................... 51 4.3.3 Modal Analysis ..................................................................................................... 52
Honeycomb Solar Panel Wings ............................................................................................. 54 55.1 Requirements .................................................................................................................... 55
5.2 Proposed Design Concept ................................................................................................. 57
5.3 Failure Modes ................................................................................................................... 58
5.3.1 Sandwich Panel Failure ......................................................................................... 58 5.3.2 Insert Failure ......................................................................................................... 61
5.4 Honeycomb Composite Panel Design .............................................................................. 65
5.4.1 Panel Procurement ................................................................................................ 67
Deployable Components ........................................................................................................ 68 66.1 VHF Antennas .................................................................................................................. 68
6.1.1 Requirements ........................................................................................................ 69
6.1.2 Research ................................................................................................................ 69
6.1.3 Design ................................................................................................................... 73 6.2 Langmuir Probes ............................................................................................................... 78
Ground Support Equipment ................................................................................................. 79 77.1 Assembly and Handling GSE ........................................................................................... 79
7.2 Protective Enclosure ......................................................................................................... 82 7.3 Mock-Up Wings ................................................................................................................ 83
7.4 Radio Frequency Testing GSE .......................................................................................... 84 7.5 Deployment and XPOD Loading GSE ............................................................................. 86
Integration and Testing ......................................................................................................... 88 88.1 Structural Fit Checks and Preparation .............................................................................. 88 8.2 Wiring Harness Fit Check ................................................................................................. 89
8.3 Dirty-Sat Integration and Testing ..................................................................................... 90 8.4 Dirty-Sat EMC Testing ..................................................................................................... 92
8.5 Antenna Pattern Testing .................................................................................................... 93 8.5.1 Uplink/Downlink and GPS Antennas ................................................................... 93
8.5.2 VHF Antennas ...................................................................................................... 94
8.6 Deployment Testing .......................................................................................................... 94 8.6.1 Results ................................................................................................................... 95
8.7 Structural Bake-Out .......................................................................................................... 97 8.8 Flight Integration .............................................................................................................. 99
Conclusions ........................................................................................................................... 100 9References .................................................................................................................................. 101 Appendix A ................................................................................................................................ 103
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Table 1: SFL satellite platform specifications [1] ........................................................................... 2 Table 2: Driving requirements that affect the structural design [8] [9] ........................................ 16 Table 3: Component layout constraints for NORSAT-1 .............................................................. 22 Table 4: Stress and displacement analysis results summary ......................................................... 51 Table 5: NORSAT-1 honeycomb panel design failure modes summary ..................................... 66
Table 6: List of main manufacturing defects found through inspection and fit checks ................ 88
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Figure 1: The NORSAT-1 mission patch [1] .................................................................................. 3 Figure 2: NORSAT-1 microsatellite with overall dimensions ....................................................... 5 Figure 3: CLARA payload mechanical design ............................................................................. 11 Figure 4: Langmuir Probe payload mechanical design ................................................................. 13 Figure 5: Norwegian coastal regions [6] ....................................................................................... 14
Figure 6: AIS Receiver payload .................................................................................................... 15 Figure 7: XPOD-Duo deployment system, Vertical mounted (A), Horizontal (B) ...................... 18 Figure 8: GNB bus AISSat-3 (left), GHGSat-D (middle), and NEMO-AM (right) ..................... 19 Figure 9: Initial NORSAT-1 structural design proposal (Scott Armitage) ................................... 20 Figure 10: Exploded view of NORSAT-1 primary structure ........................................................ 21
Figure 11: NORSAT-1 external component layout ...................................................................... 23 Figure 12: NORSAT-1 internal component layout ....................................................................... 24
Figure 13: Reaction wheel sub-assembly, CAD model (left), clean room assembly (right) ........ 24
Figure 14: NORSAT-1 panel component layouts (front/back) ..................................................... 25 Figure 15: NORSAT-1 battery pack design, Exploded (left), Assembled (right) ........................ 27 Figure 16: Division of avionics and payloads in NORSAT-1 ...................................................... 28
Figure 17: Separation plate sub-assembly design and layout ....................................................... 29 Figure 18: Worst-case hot attitudes (as viewed from the sun) ...................................................... 31
Figure 19: Worst-case cold attitudes (as viewed from the sun) .................................................... 32 Figure 20: +Y worst-case cold dimension increase due to thermal surfaces ................................ 33 Figure 21: +Z and –Z thermal surface additions ........................................................................... 33
Figure 22: CLARA integration sequence ..................................................................................... 34
Figure 23: AIS Receiver (left) and AIS Antenna (Right) accommodations ................................. 35 Figure 24: Langmuir Probe electronics (left) and Langmuir Probe cassette (Right)
accommodations ........................................................................................................................... 35
Figure 25: Wiring harness manufacturing drawing example, Payload cable ............................... 39 Figure 26: NORSAT-1 solid model wiring .................................................................................. 40
Figure 27 Panel wiring for +Y panel (left) and +X panel (right).................................................. 41 Figure 28: NORSAT-1 design evolution ...................................................................................... 43
Figure 29: Majority of the NORSAT-1 flight structural parts ...................................................... 44 Figure 30: Large attachment bracket part (left), idealized part (middle), meshed (right) ............ 46 Figure 31: FEM fastener modeling ............................................................................................... 47 Figure 32: NORSAT-1 finite element model with applied boundary conditions ......................... 48
Figure 33: Screen capture of –Z tray FEM stress results .............................................................. 50 Figure 34: Screen capture of simulation showing first mode of NORSAT-1 at 144Hz, and
subsequent mode frequency values ............................................................................................... 53
Figure 35: Relative stiffness and weight of sandwich panels compared to solid panels. Note that
the numbers shown are normalized to the solid material numbers [13] ....................................... 54 Figure 36: NEMO-AM (A) and NEMO-HD (B) honeycomb panels ........................................... 55 Figure 37: NORSAT-1 solar array concept design ....................................................................... 57 Figure 38: Wing attachment bracket design ................................................................................. 58
Figure 39: Sandwich panel failure modes [15] ............................................................................. 59 Figure 40: Simplified loading scenario for sandwich panel failure mode calculations [14] ........ 61 Figure 41: Insert loading scenarios [18] ....................................................................................... 62
Figure 42: Partially potted insert under tensile load [18] ............................................................. 63
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Figure 43: Face sheet in-plane insert failure modes [18] .............................................................. 64 Figure 44: NORSAT-1 solar panel wing design ........................................................................... 65 Figure 45: AISSat-2 pre-deployed VHF antenna approximate dimensions ................................. 68 Figure 46: Tape spring bend radius .............................................................................................. 70
Figure 47: Tape spring materials explored ................................................................................... 71 Figure 48: Hold down configurations for two tape spring antennas ............................................. 72 Figure 49: NORSAT-1 antenna base exploded view .................................................................... 74 Figure 50: NORSAT-1 in the vertically mounted XPOD-Duo .................................................... 74 Figure 51: VHF antenna stowage system ..................................................................................... 76
Figure 52: NORSAT-1 side view of stowed AIS antennas on spare structure ............................. 76 Figure 53: Expected deployment volume of VHF antennas ......................................................... 77 Figure 54: Langmuir Probe cassette deployment .......................................................................... 78
Figure 55: MSGE assembly tray “legs” in various orientations of use (black) ............................ 80 Figure 56: NORSAT-1 GSE support stand ................................................................................... 81 Figure 57: NORSAT-1 GSE handle assembly .............................................................................. 82
Figure 58: NORSAT-1 protective enclosure design ..................................................................... 83 Figure 59: NORSAT-1 mock-up Wings (left), mock-up wings fitted on structure (right) ........... 84
Figure 60: RF testing GSE blocks ................................................................................................ 86 Figure 61: NORSAT-1 deployment jig design details .................................................................. 87 Figure 62: XPOD-Duo loading (XPOD-Duo GSE designed by Mike Ligori) ............................. 87
Figure 63: NORSAT-1 structural fit checks, CLARA (engineering model) installed (left), GSE
enclosure installed (right) ............................................................................................................. 89
Figure 64: Freshly built (left) and untangled (right) flight Main wiring harness ......................... 90 Figure 65: Payload wiring harness, 3D model (left), fit check in structure (middle), and untangled
flight harness (right) ...................................................................................................................... 90 Figure 66: NORSAT-1 Dirty-Sat integration ............................................................................... 91
Figure 67: NORSAT-1 EMC testing in SFL’s anechoic chamber ............................................... 93 Figure 68: Antenna pattern test setup ........................................................................................... 94 Figure 69: Deployment testing test setup ...................................................................................... 95
Figure 70: Deployment test still shot, just before antennas deploy .............................................. 96 Figure 71: Deployment with mounting plate, no contact ............................................................. 96
Figure 72: Deployment with plate, contact made ......................................................................... 97 Figure 73: NORSAT-1 flight structure bake-out setup ................................................................. 98
Figure 74: NORSAT-1 flight integration progress ....................................................................... 99 Figure 75: NORSAT-1 flight battery pack ................................................................................... 99 Figure 76: Insert strength as a function of core height for a 11mm diameter insert in a 3/16-5052-
.001 honeycomb core with aluminum face sheets under tension [19] ........................................ 103 Figure 77: Insert strength as a function of core height for a 11mm diameter insert in a 3/16-5052-
.001 honeycomb core with aluminum face sheets under compression [19] ............................... 104
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1
Microsatellites are paving the way for a new era in space technologies. Their low cost and quick
development time attracts a large new customer base, offering easier access to space. This easier
access to space promotes innovation and discovery at far greater rates than previously observed.
The Norwegian Space Center is taking full advantage of this opportunity, funding over five
microsatellites in the past ten years. While first limiting the funded satellites to maritime traffic
monitoring such as AISSat-1, AISSat-2, and AISSat-3, they now look to further leverage the
capabilities of small satellites through the NORSAT-1 mission, as part of the national space
program. NORSAT-1 is a multi-payload mission, and has three objectives: Investigating solar
radiation, space plasma research, and developing improved methods for detection and
management of ship traffic. The successful development of the NORSAT-1 platform aims to lay
the groundwork for additional microsatellites in the NORSAT series, and expand the Norwegian
presence in space and space-based ship tracking technologies. This thesis provides some insight
into the NORSAT-1 platform design, focusing heavily on the mechanical aspects of design,
analysis and testing.
1.1
The Space Flight Laboratory (SFL) is a company affiliated with the University of Toronto
Institute for Aerospace Studies (UTIAS) established in 1998, with one clear objective in mind: to
render space accessible to companies, government organizations, and end users like never
before. By lowering the high cost barrier to space, they can make way for far more opportunities,
with potential for the next generation of human presence in space. To date, SFL has developed
numerous low cost Nanosatellite (spacecraft mass <10kg) and Microsatellite (spacecraft mass
<100kg) platforms through the unique approach of training graduate students on real-customer
projects. With a “Micro-space” philosophy in mind, SFL leverages the latest technology
advances, off the shelf components, and in-house innovations in order to develop highly capable
spacecraft from conceptualization to launch with drastically reduced development time and cost
of typical spacecraft. With more than a dozen spacecraft currently in low Earth orbit (LEO), and
another nine under development or waiting for launch at the time of writing, SFL is a clear leader
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in Canada for its spacecraft capabilities and is pushing the boundaries on LEO space access. The
current director is Dr. Robert E. Zee, and at the time of writing, the company consists of
approximately 35 full-time employees, and 12 full-time graduate students. Table 1 below
outlines the specifications of some of the current SFL spacecraft platforms, and some of the
projects on-orbit and under-development.
Table 1: SFL satellite platform specifications [1]
Typical Specifications – Customization Possible – Specifications Subject to Change
CanX-2 NTS GNB NEMO NEMO-150
Spacecraft Mass 3.5 kg 6.5 kg 7 kg 15 kg up to 150kg
Spacecraft
Volume
10 x 10 x
34cm
20 x 20 x 20
cm
20 x 20 x 20
cm
20 x 20 x 40
cm
60 x 60 x 60cm
Peak Power
25ºC,BOL
2-7 W 4 -7 W 7 - 9 W 50 W up to 500W
Payload Mass 1 kg 2 kg 2 kg 6 kg up to 70 kg
Payload Volume 1,000 cm3 1,700 cm
3 1,700 cm
3 8,000 cm
3 up to 108,000
cm3
Payload Power
@ duty cycle
1-2 W
@100%
2 W @20-
30%
3 - 4 W
@100%, 6 W
max
45 W @ 40%
min, 65 W
max
50W or higher
ACS stability ~2° (1)
~5-10° ~2° (2)
~60" (3)
~2° (2)
~60" (3)
~2° (2)
~10-
20" (3)
Downlink 32 k - 1
Mbps
32 k - 1 Mbps 32 k - 2 Mbps 32 k - 2 Mbps 32k - 50Mbps
Examples CanX-2,
CanX-7
NTS AISSat-1, 2,
3, BRITE
Constellation,
EV9, CanX-
4&5
NEMO-AM,
GHGSat-D,
NORSAT-1
NEMO-HD
1. With magnetometer, sun sensor and one reaction wheel. 2. With magnetometer, fine sun sensor and
three reaction wheels. 3. With star-tracker and three reaction wheels.
1.2
NORSAT-1 is a multi-payload satellite under development at the University of Toronto Institute
for Aerospace Studies – Space Flight Laboratory (UTIAS-SFL) for the Norwegian Space Center
(NSC). The satellite will investigate solar radiation, space weather, and detect ship traffic by
means of three separate payloads: a Compact Lightweight Absolute Radiometer (CLARA), four
Langmuir Probes, and an Automatic Identification System (AIS) receiver, respectively. SFL has
been contracted to design the spacecraft platform to house these payloads into low Earth orbit.
NORSAT-1 is one of SFL’s third generation satellite platforms, the Next-Generation Earth
Monitoring and Observation (NEMO) class, which leverages the experience gained through the
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successful development of the Generic Nanosatellite Bus (GNB). It is currently under
development alongside two similar sized NEMO class missions at SFL, GHGSat-D and NEMO-
AM, who both aim to monitor levels of greenhouse gas in the Earth’s atmosphere.
Although NORSAT-1 is not NSC’s first satellite endeavor, nor is it SFL’s first time
collaborating with NSC to develop a small spacecraft (Ex. AISSat-1, 2 and 3), it does represent
Norway’s first satellite project with scientific purpose. The official mission patch for the satellite
is shown below in Figure 1, listing the main partners involved with the project.
Figure 1: The NORSAT-1 mission patch [1]
1.3
With the NORSAT-1 mission, SFL has the opportunity to extend its knowledge and experience
to designing a slightly larger spacecraft, accommodating multiple payloads. This thesis follows
the accomplishments of the author while working at SFL towards all mechanical aspect of
design, analysis and testing for the structural subsystem of NORSAT-1. Given that the
NORSAT-1 development was proposed as a two-year contract from kick-off in July 2013, this
thesis will follow the majority of the development cycle of the spacecraft. The structural design
is outlined from the mere early stages of its proposal concept, through the various design
iterations past the Preliminary Design Review (PDR), and Critical Design Review (CDR), until
the final (current) design. This design is currently undergoing flight level assembly, integration
and system level testing. As the main structural engineer for the project, the author was in charge of
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the full structural design of the spacecraft, designing the mechanical interfaces between everything
inside and on the exterior of the spacecraft, as well as the assembly, integration, and mechanical
aspects of testing on the fully built system level spacecraft. The contributions presented are vital to
the success of the NORSAT-1 mission, given the dependence of all included spacecraft components
on the ability of the spacecraft to maintain structural integrity.
The main objective of this thesis is to design the structural subsystem for the NORSAT-1
mission such that it properly accommodates all satellite subsystems, including the payloads, and
provides a safe and secure housing through all mission environments. The success of the design
will be based on its ability to satisfy the mission, system, and structural subsystem requirements.
A secondary thesis objective is to detail the novel aspects of mechanical design for NORSAT-1
to serve as a design reference for future microsatellites having similar requirements.
While the current chapter outlines a general introduction and the necessary background
information for the work to be presented, Chapter 2 gives a closer look at the NORSAT-1
mission, overviewing the three different payloads on-board, as well as the overall capabilities of
the spacecraft itself. Chapter 3 describes the author’s work on the structural design for
NORSAT-1, describing how the design was inspired through the need to meet several necessary
requirements as well as adhere to a low cost and stringent timeline. Chapter 4 illustrates how the
structural design was validated analytically through the use of finite element analysis (FEA)
software, in order to ensure the design would stand up to the harsh environments presented
during a rocket launch. Chapter 5 details the design of NORSAT-1’s large solar array, through
the use of two honeycomb composite sandwich panels. Chapter 6 presents the design and testing
of two deployable antennas aboard the spacecraft to receive very high frequency (VHF) signals
for the AIS receiver payload. Chapter 7 then outlines the mechanical ground support equipment
(MGSE) that was designed in order to facilitate the assembly, integration, and testing (AIT) of
the spacecraft, and Chapter 8 depicts the integration and testing of the spacecraft that the author
was directly involved with, including the dirty and clean assembly, antenna pattern testing, and
deployment testing. Lastly, Chapter 9 presents concluding remarks on the experience through the
development of the structural subsystem for NORSAT-1, and outlines the current status of the
project.
5
2
This section provides a closer look at the NORSAT-1 mission, detailing the three payloads on
board, as well as the spacecraft capabilities. Figure 2 displays an external view of NORSAT-1
with approximate outer dimensions of the satellite bus.
Figure 2: NORSAT-1 microsatellite with overall dimensions
2.1
The NORSAT-1 satellite is currently slated to launch in the first quarter of 2016 into a dawn-
dusk orbit. It is required that the spacecraft remain fully functional for at least one year in orbit,
however, the design goal reaches for at least three years. The mission will allow for simultaneous
operation of all three payloads on-board, and will be operated from a ground station in Norway.
The performance of the satellite is designed to meet the specific requirements of the on-board
payloads set forth by each of the payload providers, as well as the Norwegian Space Center.
2.2
The NORSAT-1 spacecraft is comprised of numerous subsystems that are all vital to the
spacecraft performance. Each of these subsystems for the NORSAT-1 mission is described in
brief below, in order to provide a full overview of the spacecraft design and capabilities.
6
2.2.1
The structural subsystem comprises of all of the physical components needed to properly house
the satellite avionics, payloads, and any other necessary equipment, and keep them safe through
all expected environments. It is comprised of a heritage design concept from the GNB, utilizing
two intricately designed loadbearing aluminum trays, housing much of the avionics, enclosed by
six aluminum body panels. The structure of the satellite is the main topic of this thesis and is
more thoroughly detailed in Chapter 3.
The total estimated mass of the satellite (at the time of writing) is approximately 16kg. When
loaded inside its separation system, it has a total launch mass under 30kg, and an approximate
launch volume of 300mm x 200mm x 500mm.
2.2.2
The Telemetry and Command subsystem on NORSAT-1 provides a full-duplex, bi-directional
radio communications system between the satellite and the Earth station. It incorporates a two-
radio system: a UHF receiver for uplink communications, and an S-Band transmitter for
downlink. The uplink UHF receiver is supplemented with a cavity band-pass filter to provide
rejection of the spacecraft’s transmitter emissions, a down-converter that provides frequency
translation between S-band and ultra-high frequencies (UHF), and a UHF receiver that includes
demodulation, bit synchronization, and descrambling functionality [2].
Both of these radios are heritage designs used on previous SFL missions, and utilize a dual-patch
antenna system, with the antennas for each link located on opposite sides of the spacecraft; a
total of four S-Band patch antennas are used for this, and provides close-to omnidirectional
coverage.
2.2.3
A fully passive thermal design for NORSAT-1 has been designed by the thermal engineer for the
majority of the considered orbits. The techniques used involve controlling the overall bus
temperature using various thermal control tapes on the outside of the spacecraft, as well as
controlling the internal component temperatures through materials selection and specific
mounting methods. This type of implementation allows for a low-risk and robust satellite thermal
design. In addition to these passive techniques used, an active heater is present in the battery
pack in order to ensure the battery cells are kept above their minimum charging temperature
(0°C). The thermal design will be updated once the orbit of NORSAT-1 is confirmed.
7
2.2.4
Attitude and control of the satellite, while in orbit, is achieved through various spacecraft
mounted sensors and attitude hardware. Six SFL-designed sun sensors are present on the
spacecraft pointing in the six principle directions for attitude determination, along with a three-
axis magnetometer to determine the local magnetic field. A three-axis rate-sensor is also
incorporated to provide additional attitude information when in eclipse. Three-axis control of the
satellite is achieved using three orthogonally mounted reaction wheels and magnetorquers. A
dedicated on-board computer is present on the satellite to operate the necessary attitude
algorithms. The satellite with the above mentioned attitude hardware, is capable of achieving
pointing within +/-5°. However, with the addition of a precision sun sensor mounted directly on
the CLARA instrument, NORSAT-1 can achieve a fine pointing mode, enabling pointing within
+/- 0.5 degrees (mean plus 3-sigma) while the sun is visible [2]. Additionally, a GPS receiver is
included to provide positioning and timing data, as well as support the payload activities.
2.2.5
Three identical SFL designed on-board computers comprise the Command and Data Handling
(C&DH) subsystem on NORSAT-1. These on-board computers are heritage designs from
previous SFL missions, and represent highly mature hardware and associated software for
controlling all of the spacecraft functions, as well as communications.
The Housekeeping Computer (HKC) is typically dedicated to performing housekeeping tasks on
the spacecraft, such as collecting regular telemetry from each component. The Attitude
Determination and Control Computer (ADCC) is typically dedicated to performing attitude and
determination related processing, such as reading the necessary attitude sensors and issuing
commands to the attitude actuators. The third computer on board is the Payload On-Board
Computer (POBC), dedicated to interfacing with the payloads; the Serial Interface Board (SIB) is
also considered a portion of this computer.
2.2.6
The power subsystem on NORSAT-1 has three main functions: power generation, energy
storage, and power distribution.
Power generation is done by means of externally mounted triple junction solar cells, grouped in
eight-cell strings, with a beginning of life efficiency of roughly 28%. The solar cells are mounted
on the satellite such that at least one string is visible in each of the six primary directions, to
8
allow for safe-hold power generation in any state. The main sunward facing side of the satellite
however has a total of six strings of solar cells for additional power generation during payload
activities. A total of 96 solar cells exist on NORSAT-1, capable of over 40W of power
generation.
Energy storage on NORSAT-1 is done by means of a three-series, two-parallel lithium-ion
battery pack, with an integrated Battery Interface Module (BIM). The BIM provides battery cell
protection, and provides the battery telemetry that operates the battery heater.
Power distribution on the spacecraft is achieved through an in-house developed Modular Power
System (MPS). It is comprised of a passive backplane, on which multiple Micro Switched Power
Node (uSPN) cards are connected - each load on the spacecraft is connected to an individual
switch. A 5V supply, Solar Array/Battery Regulator (SABR), Solar Array/Bus Interface Node
(SABIN), and Interface Node (IFN) card are also mounted onto this backplane to provide the
remaining functionalities to the power system.
2.3
NORSAT-1 is a collaborative mission with three separate payloads each being designed by
different companies in different countries or cities; therefore effective communication between
the spacecraft and payload design teams is essential. The use of Interface Control Documents
(ICDs) is crucial to freeze the interrelated parts of the payload and spacecraft design early on to
allow for a fully parallel design path. These documents specify and manage every interface
between the payload and the spacecraft, including the mechanical mounting interface, volume
and mass allotments. Efforts are made not to deviate from these documents in order to minimize
conflicts in schedule due to design changes. Once these documents are finalized late in the
design phase, parts can begin to be manufactured. By proceeding with the payload and spacecraft
design concurrently, the overall platform development time is significantly reduced. In this type
of approach, the spacecraft design team must be fairly involved with each of the payload designs,
in order to ensure efficient and compatible designs. Each payload has their own unique design
challenges associated with them, which often times carry over to challenges within the spacecraft
design. In the following sub-sections, the design, motivation, and goals of each of the three
payloads will be discussed, along with the key challenges that affect the spacecraft design.
9
2.3.1
The Compact Lightweight Absolute Radiometer (CLARA) instrument is the primary payload on
the NORSAT-1 mission. The CLARA payload is being designed by Physikalisch-
Meteorologisches Observatorium Davos / World Radiation Center (PMOD/WRC) in
Switzerland, and its purpose is to measure Total Solar Irradiance (TSI) with high precision and
low noise. It is a compact radiometer consisting of three digitally heat powered regulated
cavities. Each aperture gets aligned directly at the sun in order to measure TSI when the
according shutter is open. Total Solar Irradiance is one of the fundamental parameters in climate
research, and operational TSI monitors in space are crucial for climate forecast and
reconstruction [3]. The CLARA payload is being designed as compact and lightweight as
possible in order to maximize its flight opportunity on a multi-payload satellite such as
NORSAT-1. CLARA has four main science objectives that are briefly explained below:
Absolute Radiometry Validation
The CLARA payload will allow for validation of laboratory results in space that provide an
explanation for some discrepancies measured by PMO6-, DIARAD-, and ACRIM-type
radiometers compared to the American TIM/SORCE experiment [4].
Space Weather
Through successfully modeling the TSI variations, correlations to the Ultra Violet (UV) radiation
variations can also be made [2]. Thus, the long-term stability of UV variations can also be
assessed.
Climate Research
Large amounts of evidence suggest that Total Solar Irradiance (TSI) has an influence on the
Earth’s climate [5]. The CLARA payload aims to extend the TSI data record for solar
atmosphere and climate modelers through monitoring the TSI variations with great accuracy and
sensitivity. Continuous monitoring of the TSI levels is needed in order to reduce uncertainties,
and cover the 11-year solar cycle. The launch of NORSAT-1 provides a suitable timeline for
avoiding any possible gaps in data due to current missions nearing an end.
Helioseismology
The CLARA payload will be the highest-cadence radiometer in space to-date, allowing the
assessment of the TSI variability at very high frequencies. The higher frequency of
10
measurements allows for helioseismology investigations of the solar atmosphere in order to
assess the acoustic energy carried into the solar atmosphere by high frequency sound waves [6].
2.3.1.1
The measurement principle of CLARA uses black body cavities to absorb the incident solar
energy through a precise aperture, which can be closed with a shutter. The absolute irradiance
can then be measured by referencing the incident irradiance of an open cavity to the measured
irradiance of a closed cavity. This measurement technique is thermal based, and relies directly on
knowing, to a great degree of certainty, the thermal resistance of a fragile flexure structure in
each of the black body cavities. As such, all thermal variability in and around these sensitive
components will directly distort the results. Rather than trying to predict accurately the
(constantly changing) radiated and conductive thermal paths from each of the cavities in the
spacecraft on orbit, a seemingly simpler approach was taken, whereby the cavities and sensitive
components were physically separated as much as possible from the rest of the payload, in
efforts to thermally isolate them from the thermal environment. This approach led the payload
provider to the current mechanical design, as described in the following sub-section.
In addition to this thermal design challenge, the CLARA payload also requires a high level of
pointing accuracy during measurements while pointing at the sun. In order to meet the required
pointing accuracy (mean plus 3-sigma, of +/- 0.5 degrees), an additional precision sun sensor
with adequate space heritage is incorporated into the spacecraft design.
A third design challenge of the CLARA payload is the required level of cleanliness, higher than
would normally be required for even an optical telescope mission. The two areas of concern are
particle contamination inside the apertures, and hydrocarbons settling on the external thermal
control surfaces. The former is primarily a ground handling concern, whereas the latter is
primarily a spacecraft outgassing concern. To cope with this, a special handling and cleanliness
protocol has been implemented for both the payload, and the entire spacecraft. Part of the
protocol involves baking out all spacecraft components, including the other two payloads, as
well as the full spacecraft after integration, in order to reduce some of the risk of outgassing
materials later on orbit. The instrument will also have the means to be constantly purged with
nitrogen, and include a protective cover to minimize contamination.
11
2.3.1.2
The CLARA payload is separated into two separate aluminum enclosures, connected together via
four large titanium studs, and a couple of wires to transfer the measurement data. The thermally
isolated section contains the sensitive measurement components, and is wrapped with a multi-
layer insulation (MLI) blanket to further the thermal isolation.
In order to minimize any mechanical and thermal pointing misalignments between the sun sensor
and the CLARA apertures, the precision sun sensor is directly mounted on the payload via a
dedicated bracket, extending from the rear, “less thermally sensitive”, half. The solid model of
the CLARA payload is detailed below in Figure 3.
Figure 3: CLARA payload mechanical design
2.3.2
Plasmas are by far the most common phase of ordinary matter in the universe, and in our solar
system, interplanetary space is filled with the plasma of the solar wind that extends from the Sun
out to the heliopause [7]. The main purpose of the Langmuir Probe payload is to measure
electron plasma from the sun, detectable from low Earth orbit (LEO), in an effort to study and
define the plasma parameters. In LEO, at an altitude of a few hundred kilometers, the spacecraft
will primarily be submersed in the dense plasma known as the ionosphere, which is produced
from the ultraviolet radiation from the sun [7]. The instrument is being developed by the
University of Oslo (UiO), and consists of four individual probes each mounted at the end of a
12
boom (four booms total), and can assess several defining features of the plasma, such as electron
temperature, electron density, and electric potential. The system is a new concept Langmuir
Probe system capable of high-resolution measurements of space plasma density, and can cover
the density range from – . The system consists of two or more cylindrical
needle probes, and is therefore referred to as UiO’s multi-Needle Langmuir Probe system (m-
NLP). A key feature of the m-NLP technique is the ability to determine the electron density
without the need to know the spacecraft potential and electron temperature.
2.3.2.1
One of the main design challenges of the Langmuir Probe system aboard NORSAT-1 is getting
the probe tips in an area of undisturbed space plasma. As the spacecraft is in orbit, it creates a
‘plasma wake’ in the opposite direction of travel. Because this “plasma wake” is poorly
understood and hard to predict, the probe tips are placed on long protruding booms, in an effort
to place the tips as far out into the undisturbed plasma as possible. In doing this, the booms
become quite long, quickly increase in complexity, and increase the launch volume of the
spacecraft quite significantly. To minimize this, the probes must be made deployable. Having to
have a deployment system carries its own set of challenges and requirements as well, such as the
necessity of ground handling equipment and testing methods.
A second design challenge that the Langmuir Probe payload carries is the need to have sufficient
conductive surface area (coupled to the spacecraft chassis reference ground) available on each
side of the spacecraft in order to “close” the measurement circuit. As the probe tips collect high-
mobility electrons from the space plasma, the spacecraft will inevitably charge up, and could
potentially begin to repel incoming electrons if a sufficient charge is achieved. The conductive
surface area on the spacecraft sides provide a path for the less mobile ions to hit the spacecraft in
the direction of velocity, in order to offset the spacecraft charge build-up. The required surface
area is made as a spacecraft requirement, and is simply verified by inspection.
A third design challenge is that the probe tips must be free of contaminations that could impact
the measurements. For example, finger oils can create an insulating layer on the probe, and
would reduce electrons flow through the probes and affect the resulting measurements. A reliable
mitigation plan for this concern is thus required for the design, such as protective covers,
replacement tips, or repeated cleaning.
13
2.3.2.2
The Langmuir Probe tips are placed on large deployable booms, whose overall size is limited to
the length of the spacecraft itself. Two identical Langmuir Probe Cassettes are included on
NORSAT-1, each housing two Langmuir Probes. The booms are held down by a uniquely
designed mechanism that, through spring preloads, forces the booms to stay stowed. Using a
commercial shape memory alloy pin-puller, the pre-loaded spring mechanism can be released on
orbit, allowing it to perform a half-turn, and consequently push both booms out with a large
enough force to reach their fully deployed positions. Once fully deployed, each of the booms is
able to lock in place via a locking pin. This cassette design is shown below in Figure 4; also
depicted is the electronics box that accompanies the cassettes on the spacecraft.
Figure 4: Langmuir Probe payload mechanical design
2.3.3
Kongsberg Seatex, in Norway, is designing the AIS receiver payload. The AIS receiver will
detect and track maritime traffic in Norwegian and international waters via the Automatic
Identification System (AIS). The AIS system is a line-of-sight, self-organized, time division
multiple-access messaging system that provides situational awareness to a large number of
maritime vessels at sea. It allows the exchange of information such as ship identification,
position, course and speed, allowing governmental organizations to monitor and direct the ship
traffic. Mandated by the International Maritime Organization, all vessels over 300 gross tonnes
are obligated to carry the AIS system. Monitoring and collecting AIS data from space has
14
recently proven to be effective, and of great interest to the Norwegian government, due to the
large portions of open water that is currently impossible to monitor via coast-based AIS stations
(Figure 5). The icons seen near the mainland coast of Norway represent the extent of the coast-
based AIS monitoring, while the blue shaded areas represent the relevant Norwegian and
international waters.
Figure 5: Norwegian coastal regions [6]
The AIS payload consists of a dual antenna very high frequency (VHF) receiver supporting four
VHF channels each. The technology of the receiver will be similar to the previous AIS receivers
designed as the main payloads on previous SFL satellites, such as AISSat-1, which was launched
in 2010, as well as the recently launched AISSat-2. This new AIS receiver will be more
advanced and will have the opportunity to test out new detection algorithms [6]. The motivation
for ship detection via AIS is fairly clear and proven from previous AIS missions, in that
detection from space allows for a more complete picture of the activity in the waters and can
better prevent ship collisions. NORSAT-1 is also intended to add to the on-going capability of
the Norwegian government of space-based AIS systems, marking this as the fourth satellite in
15
the constellation of AIS satellites, the others being the already on orbit AISSat-1 and AISSat-2,
as well as the soon to be launched AISSat-3 – for all of which the spacecraft platform was
designed by SFL. The AIS receiver payload is the sole non-science payload aboard the
NORSAT-1 mission.
2.3.3.1
Two main challenges exist with the AIS payload. First, because of the low frequency of
operation (VHF), the antennas have to be quite large relative to the overall spacecraft size. On
previous SFL designed AIS satellites, single pre-deployed antennas were used, however, since
NORSAT-1 intends to use two orthogonal antennas, the volume that the antennas consume
would be significant. Because of this, similar to the Langmuir Probe instrument, the antennas
must be made deployable. In reducing the launch volume of the satellite, the amount of launch
vehicles able to accommodate the spacecraft increases considerably.
The second main design challenge associated with the AIS payload is that the receiver is highly
sensitive to electromagnetic radiation at its frequency of operation (VHF). At this relatively low
frequency, it is not uncommon that many electronics generate noise, and could easily interfere
with the payloads data collection. As such, a fairly strict requirement on platform generated noise
to the AIS payload is placed on the spacecraft design.
2.3.3.2
The overall dimensions of the AIS receiver payload are shown below in Figure 6. The total mass
of the instrument is approximately 1.5kg. The two accompanying deployable VHF antennas are
designed by the author, and are detailed in Chapter 6.
Figure 6: AIS Receiver payload
16
3
The main objective for this thesis is to design the structural subsystem for the NORSAT-1
mission such that it properly accommodates all satellite subsystems, including the payloads, and
provides a safe and secure housing through all mission environments. The success of the design
will be based on its ability to satisfy the mission, system, and structural subsystem requirements.
Some of these driving requirements that affect the structural design are outlined in the following
subsection.
Most of the previous, and ongoing, satellites designed by SFL leverage off of a Generic
Nanosatellite Bus (GNB) design that has proven effective through multiple missions and requires
minimal structural alterations for different projects. However, due to the larger payloads on the
NORSAT-1 mission, this design cannot be re-used, and a larger satellite design must be realized.
SFL is also in the final stages of development of a larger evolution of the GNB, the NEMO
(Next-Generation Earth Monitoring and Observation) bus. The first spacecraft to use this new
bus technology is the NEMO-AM (Aerosol Monitoring) spacecraft, which is set to launch in
2016. While catering to its own mission requirements, NORSAT-1 will make an effort to use
heritage GNB and NEMO technology wherever possible in order to reduce risk and cost.
3.1
A list of driving requirements that largely affect the structural design of NORSAT-1 is detailed
below in Table 2. These requirements are compiled from various sources, including overall
mission programmatic desires, payload specific constraints, lessons learned from previous SFL
missions on the design, assembly, and testing, as well as requirements directly from the launch
vehicles. The success of the structural design of NORSAT-1 hinges on its ability to satisfy the
listed requirements.
Table 2: Driving requirements that affect the structural design [8] [9]
# Requirement Comments/Rationale
General Requirements
1
The structure shall support and ensure survival of all spacecraft components through integration, transport, handling, and launch.
Basic definition of the structural system. Cannot be verified until post-launch.
17
2 The mission should use GNB and/or NEMO heritage components to the extent practical.
Programmatic desire to minimize Non Recurring Engineering (NRE) and cost in general.
3 The spacecraft dimensions, including appendages, shall be compatible with a qualified SFL satellite deployment system.
In order to leverage an already existing SFL deployment system, such as the XPOD-Duo.
4 The spacecraft mass shall be less than 20 kg. Maximum spacecraft mass supported by the XPOD-Duo.
5
The subsystems and components used in the construction of the spacecraft shall be composed of materials that exhibit a total mass loss of no more than 1% of the component’s initial mass, and that contain no more than 0.1% collected volatile condensable material.
Desire to prevent material degradation and minimize deposit build-up on sensitive surfaces.
6
To allow air to easily escape from the satellite during launch, all volumes containing air shall be vented using an aperture with area (mm2) no smaller than 7x10-6(mm-1) x V, where V (mm3) is the volume of air.
Desire to prevent stresses induced by pressure differentials in vacuum.
Payload Requirements
7
The CLARA payload shall be accommodated such that its sensor apertures see the sun during nominal operations, and such that its aperture is the furthest protruding face of the spacecraft in its line of sight direction.
Needed in order to make solar measurements. Avoids any significant thermal impact from other components during measurements.
8 The Langmuir Probes (qty. 4) shall be accommodated externally parallel to each other and orthogonal to the CLARA Line of Sight.
Need to be as far away from satellite as possible, in order to be more submersed in the plasma environment.
9
The AIS antennas (qty. 2) shall be accommodated externally and be pointed orthogonal to each other and orthogonal to the Langmuir Probe booms.
In order to utilize polarization discrimination to improve the AIS message detection rate.
10
The platform shall limit platform-generated noise at the input to the AIS payload to -124 dBm measured in a 25 kHz bandwidth, within the band 156.025-162.025 MHz.
As measured by the AIS payload. Desire to limit platform noise propagation to the AIS receiver so as to not affect payload measurements
Launch Vehicle Requirements
11
The spacecraft must be capable of surviving, with positive safety margins, expected quasi-static launch loads of the launch vehicles under consideration, which is defined as the 5-sigma acceleration value of the composite random vibration spectrum.
In order to ensure the spacecraft will survive all launch loads without yielding. The 5-sigma value is used to instill a high level of confidence in the design. Due to launch uncertainty, all launch vehicles are considered.
12
All spacecraft components must have a first natural frequency (FNF) in excess of that required by the launch vehicles under consideration.
In order to prevent dynamic coupling between the spacecraft and launch vehicle. Due to launch uncertainty, all launch vehicles are considered. The PSLV imposes the most severe requirement, where the spacecraft must have a FNF greater than 90Hz.
18
13
In its flight configuration the satellite shall be subjected to an acceptance-level vibration test at levels specified by the launch provider and must pass this test without failure.
Required by the launch provider to prove the spacecraft can handle the expected loads.
Assembly and Testing Requirements
14 All mechanisms shall be testable in a 1g environment, with suitable GSE.
To verify their functionality in a more stringent environment.
15 Threaded inserts (Helicoils) shall be used in all non-steel materials requiring threading (aluminum, magnesium, etc.) where possible.
Prevents thread damage to expensive custom components.
16 The use of nuts should be avoided in favor of mounting bosses with threaded holes.
Reduces the number of components in the spacecraft and reduces risk of small components coming lose during launch. Also simplifies the assembly/disassembly process.
17 The spacecraft structure should allow access to any subsystem component without requiring full system disassembly.
Desire to simplify the assembly/disassembly process and to enable rapid de-bugging.
18 The spacecraft should not require custom tools to assemble or disassemble.
Custom tooling is expensive and its necessity complicates integration and testing activities, especially if done off-site.
A major defining constraint for the overall geometry and mass of NORSAT-1 is the desire to use
an existing SFL designed deployment system (Requirement #3 and #4). This greatly reduces
costs associated with Non-Recurring Engineering (NRE) because the design of the spacecraft can
then leverage large amounts of design work from previous projects. The XPOD-Duo is presented
below in Figure 7; its design is largely based off the successful XPOD deployment system that
was developed for the GNB class of nanosatellites at SFL, and it was chosen for the NORSAT-1
mission due to its larger capacity.
Figure 7: XPOD-Duo deployment system, Vertical mounted (A), Horizontal (B)
The XPOD-Duo houses the satellite aboard the launch vehicle into orbit, and upon command,
ejects the satellite by means of a compressed spring. Four points of contact are required by the
spacecraft to mate with the deployment system (known as the satellite “feet”), as well as four
19
launch rails to provide a smooth ejection. The XPOD-Duo also features two open faces to
accommodate spacecraft external appendages such as antennas and solar panels. During the
course of this thesis, the XPOD-Duo has been slightly redesigned from Figure 7A to Figure 7B –
with the main change being that it will be horizontally mounted on the launch vehicle instead of
vertically. Due to this change midway during the design of NORSAT-1, some minor changes
were made to ease the accommodation and will be discussed later.
3.2
The design builds on the design methodology of the GNB nanosatellite, whereby two trays are
used to mount large and/or massive components (e.g. reaction wheels, batteries, radios, etc.). The
satellite is then enclosed using metallic panels onto which solar cells or other light
deployable/pre-deployed components can be attached; these panels also provide a degree of
additional structural support. Each panel is 2mm thick with additional cross braces machined
directly onto the inward facing surface to increase panel natural frequencies, reduce deflections
under acceleration and increase stiffness during machining. The trays are positioned on opposing
sides of the satellite, leaving a relatively large volume between them at the center of the satellite
available for the internal payloads. This concept has been successfully implemented on the GNB
bus for multiple missions, and has been extended to the NEMO bus; therefore, NORSAT-1 will
continue to implement the proven design concept. This concept can be seen below in Figure 8,
showing the 3D solid model for various GNB and NEMO bus designs.
Figure 8: GNB bus AISSat-3 (left), GHGSat-D (middle), and NEMO-AM (right)
3.2.1
Upon the authors joining of the NORSAT-1 project in October 2013, SFL had just undergone the
Preliminary Design Review (PDR) for the project; placing NORSAT-1 in the detailed design
phase of development. Two fellow SFL workers, Scott Armitage and Jamie Fine, had worked on
20
the preliminary structural design prior to the PDR, including generating an initial structural
design proposal of the satellite bus, shown below in Figure 9.
Figure 9: Initial NORSAT-1 structural design proposal (Scott Armitage)
Initially, large volumes were allotted for the payloads due to the uncertainty of their designs,
which largely drove the size of the concept at this initial stage. These volumes then significantly
decreased as the payload designs progressed. One large solar array was included to generate the
required power during operations while the CLARA payload is directed at the sun.
3.2.2
The primary structure of NORSAT-1 (Figure 10) consists of a pair of aluminum trays (+Z and –
Z) on which most of the avionics and payloads are attached, a set of panels (+X, -X, +Y, -Y, +Z,
-Z) and risers (+X and –X) that form the outer structure, and an internal separation plate that
separates the avionics from the payloads. The risers serve to extend the bus volume through one
of the openings on the XPOD-Duo for added capacity. The rails and feet that interface with the
XPOD-Duo are integral to the trays, which form the main load-bearing structure. A separate
dedicated bracket for the three reaction wheels is included, and attaches directly to the –Z tray,
along with a number of attachment brackets for the large solar array wings.
The entire primary structure is machined out of aluminum 6061-T6, commonly used in aerospace
applications, in order to obtain the required tolerances and mechanical properties. Numerous M3
sized stainless steel screws are used to fasten everything together, with the majority of the
threaded hole features machined directly into the two trays, in compliance with Requirement #16
in Table 2.
21
Figure 10: Exploded view of NORSAT-1 primary structure
Table 3 below lists all of the main components that must be positioned in the NORSAT-1
structure. In it, the table briefly lists the main constraint or subjective requirement regarding its
placement in the satellite, in addition to those specifically highlighted in Table 2. For each of the
components, some general guidelines that helped shape the overall layout and were implemented
where possible include the following:
- Position heavier components towards the geometric center of the bus.
- Position antennas to allow for omni-directional coverage.
- Minimize the number of electronic boards in a single stack.
- Avoid mounting any components on the solar panel wings besides solar cells.
- Minimize wiring unrelated to the payloads in the payload volume.
22
Table 3: Component layout constraints for NORSAT-1
Component Layout Constraint or Requirement
Solar Panel Wings Must be positioned such that they do not interfere with integration into the XPOD, and must face in the same direction as the CLARA apertures.
Reaction Wheels (3) Must be aligned with the principle body axes and mounted orthogonal to each other.
Sun Sensors (6) Must be able to see in all six principle directions.
Magnetorquers (3) Must be aligned with the principle body axes and mounted orthogonal to each other.
Magnetometer Should be mounted as far away as possible from the reaction wheels, large current sources, and large magnetic dipoles. Must be aligned with the principle body axes.
Rate Sensors Must be aligned with the principle body axes.
House-Keeping Computer (HKC)
Locate as close as possible to the ADCC to minimize wiring.
Attitude Determination and Control Computer (ADCC)
Locate as close as possible to the HKC to minimize wiring.
Payload On-Board Computer (POBC)
Locate as close as possible to the HKC and SIB to minimize wiring.
Power Avionics/Modular Power System (MPS)
Locate as close as possible to the batteries to minimize wiring.
Batteries All cells should be located as close together as possible and as close to the power avionics as possible to minimize wiring.
Serial Interface Board (SIB) Should be placed close to the payload connectors and POBC to minimize wiring.
S-band Transmitter Should be enclosed in a similar fashion to GNB with coaxial connections near +Y end of spacecraft for ease of access during assembly.
S-band Down-converter Should be close to the UHF receiver to minimize wiring.
S-band Combiner Should be close to the S-band Cavity Filter to minimize wiring.
S-band Cavity Filter Should be close to the S-band Down-converter and Combiner to minimize wiring.
UHF Receiver Should be enclosed in a similar fashion to GNB.
GPS Receiver Should be mounted near the GPS Antenna to minimize wiring.
GPS Antenna Must be aligned such that the Solar Array and main body do not interfere with coverage. Should not be mounted to the solar panel wings.
S-band Patch Antennas Must be aligned such that the Solar Panel Wings and main body do not interfere with omni-directional coverage. Should not be mounted to the solar panel wings.
CLARA Payload Must be pointed out the main sun-facing side.
Precision Sun Sensor Should be attached to the same mounting structure as CLARA to minimize thermo-elastic distortions between them.
AIS Receiver Payload Should be as close as possible to the VHF antennas to minimize wiring, and should be in a RF noise reduced area.
AIS antennas Should be as close as possible to the AIS receiver to minimize wiring. Must be outside the spacecraft “Faraday cage”.
Langmuir Probe Electronics Should be positioned as close as possible to the Langmuir probe cassettes to minimize wiring.
Langmuir Probe Cassettes Positioned externally, with connectors as close as possible to the Langmuir Probe electronics to minimize wiring.
23
In addition to these component layout constraints, a number of external surface area
requirements also exist that reserve amounts of area on each side of the spacecraft. For example,
an area sufficient for one solar cell string of eight cells shall be reserved on the structure in each
six directions.
From the above-mentioned constraints and requirements, a component layout for NORSAT-1
could be realized. The final external and internal component layout is depicted in Figure 11 and
Figure 12 respectively. Much of the internal layout design of the spacecraft electronics was
leveraged from another SFL on-going project of similar size and requirements (GHGSat-D) in
order to minimize costs and time due to Non-Recurring Engineering (NRE) during the design
phase. The payload volume used in this satellite is large enough to house the three internal
payloads for NORSAT-1; therefore much of the avionics component layout could be left as is.
Figure 11: NORSAT-1 external component layout
24
Figure 12: NORSAT-1 internal component layout
3.2.2.1
The –Z tray sub-assembly houses all three on-board-computers (HKC, ADCC, POBC), the rate
sensor, the –Z sun sensor, the S-band Transmitter, and the UHF Receiver. The three computer
boards are stacked together using aluminum spacers, and are recessed into the tray via a
machined-in housing. The rate sensor has its own housing to mount the sensors in each axis, and
the housing is directly mounted on the –Z side of the tray; on this same side is mounted one of
the sun sensor boards for the –Z panel. The radio enclosure on the +Z side of the tray
incorporates a nearly identical housing for the S-band transmitter and UHF receiver to that which
is used on the GNB bus. This was done in order to keep a heritage design and avoid potential
compatibility issues between the already designed receiver/transmitter boards and its mechanical
enclosure.
A dedicated reaction wheel bracket is designed to house all three reaction wheels in their
required orientation, in order to provide a compact and modular integration of the wheels into the
satellite. This bracket is mounted as a sub-assembly onto the –Z tray, and it is shown below in
Figure 13. In order to facilitate the wheel integration with the wiring harness, the wheels are not
included at the sub-assembly level of the –Z tray, but are integrated as their own sub-assembly.
A similar approach is taken with the CLARA payload.
Figure 13: Reaction wheel sub-assembly, CAD model (left), clean room assembly (right)
The avionics housed on the +Z tray sub-assembly include the Modular Power System (MPS),
battery pack, GPS receiver, S-band down-converter, S-band combiner, S-band cavity filter, as
well as the AIS receiver and Langmuir probe electronics payloads on the lower half. The +X side
of the tray is slightly recessed in order to compactly accommodate some of the taller components
such as the modular power system and battery pack. These components are directly fastened onto
the tray via machined-in features at the mounting points, which also raise them off the surface in
25
order to provide clearance for the board mounted components, while also providing an area to
strategically place thermal control materials as required by the thermal design.
Some of the lighter components, such as the sun sensors and magnetorquers, are directly
mounted on the back of each of the panels, as well as all of the antennas and solar cells on the
front. Each panel is fitted with a panel connector for its sub-assembled components, which then
mates to the main wiring harness of the bus when the panel is integrated. Each of the panels’
sub-assemblies is depicted below in Figure 14. Note that there are some multiples of similar
components on different panels (six sun sensors, three magnetorquers, etc.) and have only been
labeled once in the figure for simplicity. The different colors seen on the front of some of the
panels represent a proposed thermal tape scheme based on the thermal design for a certain orbit.
Each panel is also fitted with a temperature sensor, which is not shown in the figure.
Figure 14: NORSAT-1 panel component layouts (front/back)
As seen in the above figure for the +Z, +X, and –X panels, the spacecraft avionics are kept in the
top (+Y) half of the panel, in order to keep these components in the avionics bay of the satellite.
26
The only non-payload components in the payload bay of the satellite are the –Y sun sensor,
temperature sensor, solar cells, and uplink/downlink antennas. Polycarbonate screws are used to
mount the uplink/downlink antennas and the magnetometer in order to not affect their
performance. A single string of eight solar cells are included on the +Y, -Y, +X, and –X panels.
For the +Z and –Z sides of the spacecraft, solar cells are placed on the front and back of the solar
array wings. For the Y panels, the solar cells are directly laid on the aluminum substrate, and
alternatively on the X panels, they are laid on a separate aluminum coupon, and then mounted on
the X panel over stainless steel spacers. The outer appendages, such as the AIS antennas,
Langmuir probe cassettes, and solar panel wings are designed to be installed after the entire
spacecraft is integrated, via external mounting points and panel mounted skin connectors.
The battery pack design incorporates six SAFT rechargeable lithium ion battery cells, two
parallel strings of three cells in series. The battery cells have a much narrower temperature range
than most of the other spacecraft electronics; therefore efforts are made to control the
temperature of the battery pack separately from the spacecraft as much as possible. The cells are
arranged in a set of Delrin acetal homopolymer resin collars, which allow the heat produced by
the cells to be somewhat isolated from the spacecraft due to the materials’ low thermal
conductivity, and allows for the control of the amount of conductivity from the pack to the +Z
tray using varying amounts of thermal control materials below the pack. A heater is also needed
for each string of cells in order to keep the batteries at a safe temperature in colder scenarios;
polyimide film insulated flexible heaters are used for this. Since the Delrin material provides an
inefficient medium to spread the heat of the heater to the cells, an aluminum heater plate is
incorporated to more efficiently spread the heat – the cells are each thermally strapped together,
as well as to this heater plate via various thermal control materials (Gap Pad, Pyrolytic Graphite
Sheets (PGS)). Clearance for each cell in the collars and between adjacent cells is carefully
selected in order to allow for expected battery swelling after long-term use. The battery pack
design is displayed below in Figure 15. Note that in the flight battery pack (not shown in the
figure), the cells are thermally strapped together using PGS, and the entire exposed cell area of
the pack is wrapped in thermal tape in order to prevent thermal exchange through radiation in the
spacecraft. Temperature sensors are also included on the middle cell of each string, but are also
not shown in the figure. The fully integrated flight battery pack is shown in the later Section 8.8
Flight Integration.
27
Figure 15: NORSAT-1 battery pack design, Exploded (left), Assembled (right)
The Battery Interface Module (BIM) is directly mounted on the battery pack, but its thermal path
is separated via the BIM sink plate. The sink plate allows for the BIM to sink its thermal energy
to the spacecraft +Z tray, rather than directly to the battery cells in order to prevent thermal
gradients across the individual battery cells. A custom terminal block was also incorporated into
the battery pack design, for the sole purpose of ease of assembly. This allows for the battery cells
and BIM to be wired and prepared separately, and then assembled together in a clean room
environment. Wire tie mounts are used on the top collar for battery cell wire management.
3.2.3
As mentioned in the previous Section 2.3 Payloads, the AIS receiver payload is greatly sensitive
to Radio Frequency (RF) noise produced in and around the satellite because of its relatively low
frequency operating levels. Because of this, efforts are directly made in the structural design to
reduce the amounts of potential electromagnetic interference (EMI) propagation to each of the
payloads in order to satisfy Requirement #10 in Table 2. The strategy was to attempt to isolate all
of the noise producing spacecraft avionics from the payloads, allowing minimal RF leakage out
of the isolated avionics bay, without requiring a large redesign of the leveraged bus layout; this
was done in a number of ways:
A “separation plate” was added internally to the spacecraft, physically separating roughly
90% of the satellite electronics from the payloads, resulting in an “Avionics bay” and a
“Payload Bay” as shown in Figure 16.
EMI reducing gaskets were incorporated in the design, in an effort to render the avionics
bay close to a “Faraday Cage”, reducing RF leakage to the payloads.
28
Additional screws were used around the volume of the avionics bay to both ensure
sufficient gasket compression, as well as reduce the gap lengths between screws where
noise can potentially leak through.
EMI filtered connectors were used at any interface from the avionics bay to the payload
bay or outside the spacecraft.
Figure 16: Division of avionics and payloads in NORSAT-1
Since much of the design was leveraged from heritage projects in order to reduce development
time, there was limited time to create a dedicated design for isolating the avionics. A “best
effort” approach was taken, whereby gaskets were introduced into the design wherever they
could be easily accommodated, with minimal effect on the overall design and assembly
procedure. A gasket is placed fully around the Main and Small separation plate, on the +X and –
X sides of the +Y panel, on the +Y and –Z side of the –Z tray, on the +Z panel under the AIS
antenna mounting plate, and on the +Y panel below the GPA antenna. Other areas where it was
too difficult to introduce a gasket feature were either left as is, or the screw spacing in those
areas were decreased by means of introducing additional screws – these areas include the –Z
panel, +X and –X panel upper section on the avionics side, the +Z panel, and the risers.
Efforts were not made to render the payload bay of the spacecraft as isolated as the avionics due
to the difficulty in mounting the CLARA payload. Due to its thermal sensitivity, physically
contacting the payload anywhere other than the dedicated mounting points could skew the
scientific results - sealing something that cannot be contacted proves to be a difficult task.
Therefore, an opening is left surrounding the CLARA payload in the payload bay, and efforts are
focused on shielding the noise from escaping the avionics and potentially entering the payload
29
bay from this gap. In addition, similar low noise emitting requirements exist for each of the
payloads themselves; therefore the noise produced in the payload bay is expected to be minimal.
Filtered connectors for each of the payloads are directly mounted on the main separation plate, in
order to transfer the power and data signals from the avionics to each payload. A filtered
connector is also included on the plate for the precision sun sensor, and the –Y panel avionics
since they reside in the payload bay. The RF signal for the –Y panel uplink/downlink antennas is
syphoned through the separation plate via a pair of SubMiniature Verson A (SMA) connector
adaptors. Filtered skin connectors are also directly mounted on the –Z tray for the solar array
wings, and on the +Z panel for the Langmuir probe cassettes. The design of the separation plate
is shown below in Figure 17. The Serial Interface Board (SIB) was directly mounted on the
avionics side of the separation plate in order to facilitate its wiring, since it connects directly to
each of the payload connectors on the plate. Numerous tie mounts are epoxied to the main
separation plate for use in wire management, which will be discussed in Section 3.6 Wiring
Harness Development.
Figure 17: Separation plate sub-assembly design and layout
The separation plate is separated into a Main and Small plate in order to minimally affect the
existing tray designs. The plates assemble together via five M3 fasteners while sandwiching the
+Z tray. Together, the plates provide structural mounting points for the –Z tray, +X Panel, -X
panel, +Z panel, and both risers, while also compressing the gasket between them.
3.3
As seen during the battery pack design description in the previous section, the thermal design of
a satellite is closely linked to the structural design. Materials selections, component layout,
component mounting, and overall geometry have a large effect on the thermal results. The details
30
of each component mounting connection and layout are discussed with the thermal engineer for
NORSAT-1, Vincent Tarantini, during the detailed design phase, in order to understand the
thermal point of view of each design decision. Over the course of the project, various thermal
concerns were raised regarding the overall design and proposed working orbit, which amounted
in structural design modifications in order to better accommodate the fully passive thermal
design – these modifications are explained in this section.
3.3.1
Initially, the proposed NORSAT-1 structural design included one large solar panel to achieve the
necessary surface area for all of the solar cells during payload operations, similar to the NEMO-
AM microsatellite project (seen in Figure 9). Soon after the Preliminary Design Review (PDR),
it was discovered that this proposed large solar panel could potentially be a thermal concern and
alternative solutions were considered. The problem exists because the large solar array is able to
fully shadow the entire spacecraft bus. In the scenario where only the large solar array points at
the sun continuously, the spacecraft would need to be well thermally coupled to the array in
order to keep warm; however, if it is well coupled in other attitudes, the solar array would act as
a large radiator, and large thermal swings would be seen by the spacecraft. A solution was found
by splitting the large solar array into two separate solar array “wings”, whereby a direct heat path
to the spacecraft exist in the middle through the –Z panel for thermal control. In addition, the
solar array wings would be thermally isolated from the bus, in order to have more control over
the bus temperatures in all attitudes.
3.3.2
During the detailed design phase of NORSAT-1, the thermal simulations of the structural design
found that the satellite bus would get too hot in the “worst-case hot situations”. The worst-case
hot situation describes an orientation of the spacecraft on orbit where the average temperature is
maximized. This often occurs in an orientation where the maximum possible surface area of the
spacecraft is directed at the sun continuously, resulting in maximum heat absorption. For
NORSAT-1, this happens at the orientations seen in Figure 18.
31
Figure 18: Worst-case hot attitudes (as viewed from the sun)
Having the solar cells mounted directly on the spacecraft bus leads to a convenient mechanical
design, however each solar cell then acts as a large heat absorber into the bus. Due to the number
of solar cells in view in the worst-case hot orientations, the incoming heat produced becomes too
significant for the passive thermal design to regulate. The solution to this problem was to isolate
some of the solar cells from the bus by mounting them off the spacecraft, in order to achieve the
correct balance of incoming heat. This was done in two ways: the solar cells originally mounted
on the +Z panel were moved to the rear of the solar panel wings, and the solar cells on the X
panels were moved onto dedicated aluminum coupons, which are mounted to the X panels with
stainless steel spacers at the mounting points. The solar cells on the Y panels were left directly
mounted on the spacecraft panels in order to maintain a thermally safe worst-case cold situation,
described below.
3.3.3
In addition to getting too hot, the thermal simulations were also finding that the satellite would
get too cold in the “worst-case cold situations”. The worst-case cold situation describes an
orientation of the spacecraft on orbit where the average temperature is minimized. This often
occurs in an orientation where the minimum possible surface area of the spacecraft is directed at
the sun continuously, resulting in minimum heat absorption. For NORSAT-1, this happens at the
orientations seen in Figure 19.
32
Figure 19: Worst-case cold attitudes (as viewed from the sun)
Due to the numerous components mounted on the Y panels (GPS antenna, eight solar cells, two
s-band patch antennas, sun sensor, antenna hold down hitch), there is minimal room for thermal
control tapes that would allow additional heat from the sun to be absorbed. Because of this, the
amount of heat able to be absorbed in these orientations was not sufficient to maintain the
spacecraft within the temperature limits. After exploring several options, including relocating
some of the Y panel mounted components and imposing on-orbit attitude constraints, it was
decided that the most simple and minimal impact solution would be to add additional surface
area to the satellite in these cold attitudes. This was done by extending the –Y panel, and adding
a pair of flat aluminum “thermal surfaces” to the mid-plane of the satellite, which would
essentially extend the useable thermal surface area of the satellite in the worst-case cold
orientations by about 180%. By mounting these on the mid-plane of the satellite, there is an
added benefit of having the incoming energy heat up the satellite more effectively from the
middle of the spacecraft, rather than from one end. Another large benefit to this solution is that
while a large surface area increase is seen in the worst-case cold orientations, virtually no surface
area is added in the worst-case hot orientations. The inclusion of both these additional parts
would add less than 100 grams to the mass of the spacecraft, and was deemed a low-risk, and
satisfactory thermal solution. Figure 20 and Figure 21 display the design and placement of these
added thermal surfaces.
33
Figure 20: +Y worst-case cold dimension increase due to thermal surfaces
Figure 21: +Z and –Z thermal surface additions
3.4
The way in which each of the payloads has been mechanically accommodated in the satellite in
order to satisfy the relevant requirements is discussed in this section.
3.4.1
The CLARA payload is mounted in the payload bay of the satellite, such that it is the most
protruding face in the –Z direction – this is to directly satisfy Requirement #7 in Table 2. It has a
Precision Sun Sensor mounted on a bracket that is directly attached to the rear section of the
payload in order to minimize calibration offsets through thermal and mechanical distortions. The
payload is uniquely integrated into the satellite at a late stage in order to provide minimal
handling of the sensitive payload and exposed MLI. This is done by the use of slots in the
spacecraft –Z tray, whereby the feet of the CLARA payload can slide through into the spacecraft.
Once the feet are fully inside, the payload is shifted towards the +X direction into some
34
machined-in slots, and then four M5 titanium blots can be inserted from the outside of the
spacecraft into the payload mounting feet. Finally, an aluminum side plate is installed from the -
X opening of the satellite to shield the majority of the opening left beside CLARA. This
sequence of installation is depicted in Figure 22.
Figure 22: CLARA integration sequence
Some advantages to this method of mounting CLARA are that in order to remove CLARA from
the satellite assembly after full integration, only removal of the –X panel and -X wing is needed.
This reduces much of the risk of de-integrating the satellite if removal of CLARA is needed at a
late stage. It also allows CLARA to be installed at a very late stage in the assembly procedure –
minimizing potential contaminations through handling.
3.4.2
The AIS receiver is mounted on a set of structural cross braces machined into the +Z tray using
eight M4 fasteners and is located in the noise-reduced payload bay; connector access is achieved
through the –X side of the spacecraft. Both of the VHF antennas are placed on the +Z panel,
orthogonal to each other (Requirement # 9 in Table 2), and 45 degrees to the satellite face. These
accommodations are shown in Figure 23, with the relevant mounting holes highlighted in red.
35
Figure 23: AIS Receiver (left) and AIS Antenna (Right) accommodations
3.4.3
The Langmuir Probe electronics are mounted on the opposite side of the cross braces supporting
the AIS receiver on the +Z tray using four M3 fasteners. The two cassettes of probes themselves
are mounted in parallel on either Riser using eight M3 screws each (Requirement #8 in Table 2).
They connect to the Langmuir probe electronics box by means of an intermediate wiring harness,
which allows each cassette to be plugged directly into the +Z panel from the outside. These
accommodations are shown in Figure 24, with the relevant mounting holes highlighted in red.
Figure 24: Langmuir Probe electronics (left) and Langmuir Probe cassette (Right)
accommodations
3.5
Considerations are made during the design to ensure assembly and disassembly is both possible,
and convenient (Requirement #17 in Table 2). The design is kept modular to allow a fairly
smooth integration process. Each sub-assembly is first fully integrated individually, and then all
of the sub-assemblies are integrated together. The large wiring harnesses are first integrated to
the +Z tray sub-assembly, followed by the joining of both tray assemblies via the separation plate
and some ground support equipment (GSE); then begins panel integration. Due to the loss of
access once panel assembly onto the trays begins, a specific order is derived for the assembly
36
procedure to ensure that the necessary access is available at all stages of integration. For
example, access is needed from the –Y end of the satellite in order to mate the Langmuir probe
cassette connectors to the Langmuir probe electronics box, therefore, the +Z panel must be
integrated prior to the –Y panel. The full detailed procedure for the sub-assembly and system
level assembly and integration for NORSAT-1 has been developed by the author in [10], and has
been summarized at a high level in the list below.
1) Integrate all sub-assemblies
2) Install the large wiring harness on the +Z tray
3) Connect the –Z tray to the +Z tray sub-assembly
4) Connect and route all of the wiring harnesses
5) Integrate the +Z panel sub-assembly
6) Integrate the –Y panel sub-assembly
7) Integrate the Reaction Wheel sub-assembly
8) Integrate the +Y panel sub-assembly
9) Integrate the +X panel sub-assembly
10) Install the +X and –X riser
11) Integrate the –Z panel sub-assembly
12) Integrate the CLARA payload
13) Integrate the –X panel assembly
14) Integrate both Langmuir Probe Cassettes
15) Install both AIS antennas in stowed position
16) Integrate both solar panel wings
The connectors on most electronics are placed such that connector access is always achieved
from the X sides of the spacecraft, allowing most connections to be made during panel
integration from the large open sides, and renders the X panels to be the last panels integrated.
The main bulkheads from the Modular Power System, CLARA, the AIS receiver, the separation
plate connectors, the UHF receiver, the S-band transmitter, the SIB, and the OBC connectors are
all accessible from the +X or –X side of the spacecraft during integration. This also allows for a
large amount of debugging to be done during the testing phase with the removal of just one panel
of the spacecraft.
Skin connectors are incorporated for the solar array wings and Langmuir probe cassettes on the
–Z tray and +Z panel respectively. This allows for these large protruding elements to be installed
at the final stage of integration to simplify handling during the rest of the assembly. The same
strategy is applied to the AIS antennas, which are designed to be mounted from the outside, and
can thus be integrated last. The AIS antenna mounting plate is designed as a separate piece from
the +Z panel for ease of assembly as well. This allows for the antenna connectors to be soldered
to the gold plated antenna mounts as their own sub-assembly, and then mounted to the +Z panel
sub-assembly separately.
37
The proposed solar cell layout on the solar array wings minimizes the size of the panels; however
it forces the panels to have a larger Y dimension than the satellite itself, which can cause certain
assembly complications. To accommodate this, the wings are purposefully not centered about the
Y-axis of the satellite; rather, they are offset, and biased towards the +Y end. This is done in
order to allow for the fully integrated satellite to rest on the –Y satellite “feet”, and ease the
integration of the wings in this orientation, and not limit the potential resting attitudes of the fully
integrated satellite. This was favored over enlarging the panels in the X direction, which would
force the center of mass of the cantilevered panels further outward.
3.6
While designing the spacecraft layout and structure, careful consideration must be made for the
planning of wire management. The structural design can easily incorporate simple inclusions to
aid wire routing and management, through means of tie down points and wire path cutouts, and
thus should not be left to a late stage in the design. A clean, robust wiring harness in the fully
assembled satellite is sought, in order to ease the burden of accessing components after
integrations, and to avoid loose wire bundles shaking violently during launch vibrations,
potentially de-mating connections. Some general guidelines that were followed for the wire
harness development are explained below.
Design for Assembly – The harness layout should be optimized for assembly efficiency during
integration.
Minimize lengths of wires –To save mass, to minimize signal power loss along the wires, and to
avoid turning them into Radio Frequency (RF) antennas or current loops (magnetic fields).
Minimize connectors – Based on the already designed hardware, namely the Modular Power
system (MPS), certain signal levels are limited to specific outputs from the MPS, which then
may connect to several different components, causing large, interconnected, harnesses with
numerous connectors. These large harnesses are difficult to build, manage, and integrate.
Clean routing – Common tie down points should be used for multiple components, which can
simplify the assembly and integration process, minimize tie down points, and prevents the
harness from becoming a “web”, blocking access to components after integration.
Tie down locations – Wires should be tied down around every 10cm to prevent large amounts of
loose wiring and each connector should have a tie down point immediately before its mating
connector. All wires should have the necessary strain relief as well as wire minimum bending
38
radius accounted for. The free wires next to their connectors should always provide a
compressive force, so as to counteract de-mating.
The wiring harness interconnect diagram, generated by the system engineer, depicts all the
satellite electronic components and their specific pin-to-pin assignments. This is generated early
in the project with only electrical and power requirements considered for a system level
architecture. This was used to determine all wire paths in the satellite for initial wire routing and
component layout design. This diagram was then converted into graphical harness manufacturing
drawings for each individual harness on NORSAT-1. Including each panel harness, this
amounted to 23 individual wiring harnesses with over 100 total connectors, each harness ranging
from having just one connector, to one large one having 53 connectors.
The large harness with 53 connectors posed a concern for the project because of the numerous
points of failure, and difficulty to build without error. In an effort to reduce the number of
connectors on this large harness, the author was tasked to investigate rearranging pin
assignments on some connectors. Due to the limited outputs levels on the MPS, no amount of
rearranging of pin assignments would allow the harness to be broken into smaller pieces.
However, after some pin reorganizing, a point in the harness was found where only four wires
connected two large sides. It was then decided to incorporate an “in-line” connector between
these two halves of the harness, adding two connectors overall, but splitting the large 53
connector harness into one 39 connector harness, and one 16 connector harness. The smaller of
these two harnesses was dubbed the Payload (and sun sensor) harness, connecting the power and
data lines to all of the payloads and sun sensors, while the larger one was dubbed the Main
harness, connecting all the remaining spacecraft avionics.
Each graphical manufacturing drawing, created by Payam Mehradnia, was then supplemented by
the author with additional information and instruction for the build, such as specific wire lengths,
colors, gauge, twisted pairs, splice orientations, connector part information, and pictures when
necessary. The SFL technicians meant to be building the physical harnesses were heavily
consulted with in order to ensure a clear and easy to understand presentation of information to
ensure a smooth harness build with minimal mistakes. This marked the first time a graphical
approach to building the wiring harness had been used at SFL. Using the Altium Designer
software, the graphical manufacturing drawings could be easily navigated digitally during the
39
build. Figure 25 below shows an example of one of the NORSAT-1 manufacturing drawings, for
the Payload cable.
Figure 25: Wiring harness manufacturing drawing example, Payload cable
The length information added to the manufacturing drawings is largely determined by the
proposed routing in the satellite, and was obtained in a number of ways, explained in the
following subsections.
3.6.1
In order to consider the wire routing at an early stage, before the satellite structure had been
ordered, the routing was virtually designed into the 3D solid model using the “Harness Design”
application in Solid Edge. All component connectors were ensured to be included in the model,
and wire paths were then introduced from connector to connector, and the specific routing could
be manipulated in 3D to the user’s choice. In order to simplify the model, bundles of wires
coming from the same connectors were modeled as a single wire of larger diameter.
Through this process, potential locations for wire tie down points were determined, and were
implemented either by including a pair of holes for a zip-tie to wrap through, or providing a
screw down point for a wire zip-tie mount or P-clip. Due to the relatively low impact of
including a pair of holes for zip-ties, additional tie down points were included in a number of
places (for example along the –Z tray where numerous wires pass) in order to allow some
flexibility on the wire routing when being implemented. The size and location of cutouts needed
in the structure in order to pass through wires could also be determined through this process. The
40
solid model internal wire routing of all the wiring harnesses is shown below in Figure 26,
including the coax cables for the radio equipment.
Figure 26: NORSAT-1 solid model wiring
Having the entire wiring harness of the satellite incorporated into the 3D solid model provides a
number of benefits to the overall development of the project. It allows for changes to the wiring
to easily be realized and implemented, it allows for a graphical depiction of each wiring path,
that can be used to facilitate and check the physical wire routing during integration, and also
provides accurate lengths and mass properties for the included wires. In order to account for
inconsistencies of modeling the wiring, such as not accounting for the presence of other wires, a
length margin is added to each of the predicted lengths, as a percent slack compensation - this
adds some margin to each length for later refinement.
The panel harnesses, being much more simple, were not incorporated into the solid model, and
the routing design was done in 2D. The routing for the +Y and +X panel is shown as examples
below in Figure 27. Tie down points are incorporated directly next to each connection point in
order to ensure a compressive force exists at the mated connector. Note that the yellow markings
indicate the use of Kapton tape, “T.S.” indicates the placement of the temperature sensor, and the
thick blue wires indicate coax cables.
41
Figure 27 Panel wiring for +Y panel (left) and +X panel (right)
Once the satellite structure was procured and arrived in house, the proposed solid model wire
routing could be validated using the physical structure and physical wires. It was found that the
lengths assumed from the solid model were quite accurate when the volume of wires in the area
was low, while the predictions in the dense wire areas tended to be slightly short. This was due to
the fact that the wiring in the solid model does not account for the physical presence of other
wires that it might have to detour in reality. However, with the added length margin to all of the
solid model estimates, this could easily be accounted for.
3.7
The main considerations for materials selection in the spacecraft include: low cost, low density,
low outgassing properties, non-magnetic, and desired strength. Some of these considerations are
alluded to in the driving requirements of Table 2, through Requirements #2, #4, and #5. The
majority of the spacecraft structure is made from aerospace grade aluminum 6061-T6. This is
commonly used in aerospace applications due to its lightweight, good thermal conductivity,
relatively high strength and low-cost characteristics. An alternate material was also considered,
magnesium ZK-31, which is used in the XPOD separation system due to its lower density and
similar characteristics to AL 6061-T6. This was however dismissed due to its higher cost, and
difficulty to machine.
Various other materials are also found in the spacecraft when different characteristics are
desired, usually for thermal purposes. Delrin acetal and G-10/FR4 are used when thermal or
42
electrical isolation is desired, or when a non-abrasive surface is needed. Delrin is used for the
battery pack top and bottom collars, and the AIS antenna guides/hitch. G-10/FR4 was chosen for
the solar array wing spacers to achieve the necessary thermal isolation, and was chosen over
Delrin due to its greater performance at higher temperatures.
All of the fasteners used in the spacecraft are Stainless Steel 316, while some of the other
stainless steel components, such as spacers for the X solar cell coupons, are 18-8 stainless steel.
Helicoil inserts are used at every mounting point in the spacecraft structure (Requirement #15 in
Table 2). They are installed in-house, and are Nitronic 60 stainless steel. These fastener and
helicoil materials selections stems from a desire to only have non-magnetic components in the
spacecraft, so as not to affect the performance of the magnetometer.
3.8
A system level mass budget was kept up to date by the author using an SFL template mass
budget used for previous missions. All components from each subsystem are included in this
budget with the latest mass measurements. When a component is still under design, or is not in
house to physically measure, estimates are used, with added contingency. For example, for all of
the structural parts, mass estimates from the solid model were used in the budget until the
structure had been manufactured and arrived in-house to physically measure. For cables,
connectors or electronic components, mass numbers are taken directly from the datasheet. As
flight components are integrated, their masses are measured and remaining contingency in the
budget is reduced to zero. Each individual screw is accounted for in the Spacer & Fasteners sheet
of the mass budget, where a detailed list of each screw, spacer, and helicoil in the assembly is
kept. This proves useful not only to keep track of the incremental mass of each of these small
components, but also in providing a full materials list of the hardware needed. At the time of
writing, just before flight integration, the mass budget tallies NORSAT-1 to be a total of
~16.1kg, which still includes ~300grams of contingency. This current mass estimate meets
Requirement #4 in Table 2. The remaining contingency is largely due to not having all of the
payloads in house to measure, and the outstanding solar array wings.
3.9
The 3D solid model of the spacecraft and all associated parts were created and managed in the
software package Solid Edge. The solid model of the spacecraft proved to be an invaluable tool
for the structural design, allowing the designer to visualize structural components, keep track of
43
the bill of materials and mass properties, and allows for seamless modification and iteration. All
the material and mass properties for each component are inputted and maintained in the solid
model in order to make use of the in-software mass and inertia calculations. Through this, a
fairly accurate estimate of the spacecraft’s center of mass and moments of inertia matrix can be
determined – which would be quite difficult to estimate analytically for something as oddly
shaped and mass distributed as NORSAT-1. These estimates are used as initial estimates in
designing the spacecraft attitude and control system, and are also used by the launch providers
when designing the satellite’s secondary payload accommodation.
A proposed thermal tape scheme for an intended orbit of the spacecraft was also incorporated
into the solid model. Having the tape scheme in the model allows for the accurate assessment of
the area coverage for each tape, and can be incorporated into the thermal model. A simple tape
scheme is desired, in order to have it easy to implement. Figure 14 of the earlier Section 3.2.2.1
shows a proposed tape scheme on each of the panels in the solid model for NORSAT-1, based on
the desired areas of coverage determined by the thermal engineer in [11]. Each different color
represents a different thermal tape with different optical properties. Note that this tape scheme is
not final, and is subject to change once the final orbit is confirmed.
3.10
The structural design for NORSAT-1 constantly evolved throughout the design process. Figure
28 below presents a look at the overall solid model at various stages of the project. The most
notable changes include the splitting of the large solar array into two solar array wings, moving
both AIS antennas to the same face, the placement of the Langmuir probe cassettes, the solar cell
accommodations, and the (not seen) inclusion of the internal separation plate.
Figure 28: NORSAT-1 design evolution
44
3.11
The majority of the structure for NORSAT-1 was manufactured by a local machine shop, with a
few small parts being machined in-house by the author. As expected, the two trays of the primary
spacecraft structure were the most expensive due to their complexity, with the remaining parts
being a fraction of their costs. Two sets of each structural component were procured, in order to
have a “flight” structure, as well as a second “spare” structure. If anything were to happen to the
flight components, a spare component could be swapped in. The spare structure is also
strategically used in parallel for various testing while the flight structure is otherwise occupied.
Figure 29 below shows the full set of the flight structure in SFL’s class 10,000 clean room.
Figure 29: Majority of the NORSAT-1 flight structural parts
45
4
In order to validate the structural design of the spacecraft, a finite element method (FEM) model
of NORSAT-1 was created for the Critical Design Review (CDR) using Siemens NX 8. The
model is continually updated post-CDR in order to reflect any major design alterations before
parts manufacturing, and to point out any unforeseen design flaws during the design process. The
analysis and post-processing of the finite element results are done using the NX Nastran solver,
and the results are then compared against the requirements to ensure a satisfactory design has
been achieved. The requirements at question are specifically #11 and #12 in Table 2, stating the
requirements set forth by the launch vehicle to ensure the design can withstand the expected
launch loads, with some added margin of safety. The details of this analysis for NORSAT-1 are
outlined in this section.
4.1
4.1.1
All of the NORSAT-1 structural components were modeled using hexahedron and tetrahedron
3D elements. Hexahedron elements were favored over tetrahedron elements for the majority of
the components for various reasons related to user setup time and computational costs.
Tetrahedron elements can be used for complex geometries and are generated automatically,
resulting in a very small user setup time. However, due to the often oddly shaped elements, quite
a large number of elements are needed for acceptable accuracy, which results in a high
computational cost, and the irregular triangular shaped elements often lead to undesirable
elements with high aspect ratios, skewing the results. Hexahedron elements are contrarily
manually meshed, and require significant effort in idealizing the part geometry for suitable
swept-meshing. Once properly meshed, the elements form a clean continuous mesh that is easily
customizable, and can provide a comparable accuracy with far fewer elements, resulting in a
smaller computational cost.
As part of the idealization process, non-essential features of each part are suppressed, such as
non-essential holes for tie wraps or components mounting, and non-load bearing rounds included
for part machinability. The remaining geometry is split into multiple sections, allowing
46
individual swept-meshes to be manually inserted. Options also exist to split the geometry at areas
of interest, and provide meshing controls, to facilitate obtaining a finer mesh at specific
locations. For example, the area around each mounting hole is split, and a mesh control is added,
to allow for a higher density of nodes to be present at these anticipated high areas of stress.
Figure 30 below shows the idealizing and meshing process for the large attachment bracket.
Figure 30: Large attachment bracket part (left), idealized part (middle), meshed (right)
Material properties are specified for each fully meshed part and applied to each created element
to ensure the correct behavior is exhibited in the simulations.
4.1.2
Due to the complexity and orthotropic material behavior of honeycomb sandwich structures used
for the solar array wings, an alternative method of modeling has to be considered. Given the
popularity in the area of modeling composites for finite element analysis, many modeling
methods exist and have been thoroughly researched and implemented for specific situations. A
study on alternative modeling methods for the honeycomb solar panel used on NEMO-AM has
been completed by Dumitru Diaconu in [12], and due to its similarity to NORSAT-1’s solar
panel wings, a similar approach to the concluded best method will be taken. The study compared
three modeling methods to some known data: an accurate three-dimension (3D) model of the
panel using a detailed model of the honeycomb core and a face sheet mesh, a two-dimension
(2D) mesh with composite shell properties, and an equivalent panel method where the sandwich
panel properties are converted to an equivalent solid panel. It was concluded that the 2D mesh
with composite shell properties method was able to produce very similar and accurate results to
that of the accurate model and equivalent panel methods, with just a fraction of the nodes and
computational effort. This was the chosen method used for modeling the NEMO-AM
honeycomb panel, and is thus the chosen method for the NORSAT-1 solar array wings.
47
The “PCOMP” composite shell implementation in the software package is used to create a 2D
mesh at the mid-plane of the part; the material properties and thickness for each layer of the
sandwich panel, the two face skins and the honeycomb core, are then inputted into the meshing
details in order to exhibit a realistic behavior of the composite. Added mass due to the adhesives,
inserts, and solar cells are added to the panel material properties as non-structural mass, which is
evenly distributed across the full geometry.
4.1.3
Much of subsystem hardware, including but not limited to, the magnetorquers, printed circuit-
boards (PCBs), the battery pack, reaction wheels, and antennas were modeled in the FEM
assembly as manually defined zero-dimensional (0D) lumped mass elements, located at the
component’s center of mass, with one-dimensional (1D) rigid element links connecting them to
their mounting locations in the spacecraft. This is done for the components that are non-load
bearing, and whose geometries would not affect the analysis results.
The payloads, contrarily, were fully modeled with 3D elements to represent their structural
design, using the latest solid model provided by the payload provider. However, in cases where
significant information was still not confirmed at the time of analysis, a similar 0D lumped mass
approach was taken to best represent the payload internal design and mechanical interface with
the satellite.
Fastener modeling was exclusively performed using infinitely stiff 1D rigid links to represent the
shank of the screw and load path between connected parts, as well as less stiff 1D rigid links to
represent the contact area of the screw head. This method has been used to model previous SFL
spacecraft structures and has been deemed to produce acceptable results. An example of this
implementation is shown in Figure 31.
Figure 31: FEM fastener modeling
48
4.2
4.2.1
The constraints for the finite element simulations aim to represent as closely as possible the
method of restraint of the spacecraft inside the XPOD-Duo separation system under launch
conditions. The ball cup interfaces at each of the eight satellite feet and the XPOD-Duo are
designed to constrain the lateral movement of the satellite, and therefore lead to fixed
translational constraints at each of the satellite feet in the X and Z directions. Since the satellite is
restricted in the Y directions when the XPOD door is closed, translational motion in the
longitudinal direction (Y axis) is fixed on the –Y feet, interfacing with the XPOD door. On the
+Y feet (located at the base of the XPOD-Duo, interfacing the with pusher plate), translational
motion is left unconstrained, in order to allow compression under longitudinal accelerations.
Finally, a compressive load is applied in the Y axis on the +Y feet in order to represent the pre-
load of the XPOD-Duo spring in the pusher plate. These boundary conditions applied to the full
FEM model are shown in Figure 32. Surface contacts were not modeled in order to minimize the
complexity of the simulations. By not including these, stress concentrations are over-estimated,
due to the lack of friction which would otherwise impede motion, leading to more conservative
results.
Figure 32: NORSAT-1 finite element model with applied boundary conditions
4.2.2
The applied loads represent the worst-case loading that NORSAT-1 will likely see while loaded
in the XPOD-Duo inside the launch vehicle. At the time of analysis, a launch vehicle had not
been confirmed; therefore all launch vehicles under consideration are taken into account to
determine the greatest acceleration likely to occur. Each launch vehicle specifies various
+Y
+X
+Z
49
vibration tests to be done on the satellite for qualification purposes prior to being accepted as a
secondary payload. Of these tests, random vibration is likely to impose the highest accelerations
to the spacecraft, and is therefore used to determine the highest likely quasi-static acceleration
that the spacecraft might see due to vibrations. This can be done by finding the standard
deviation of the accelerations experienced by the satellite if exposed to the composite random
vibration spectrum, which includes the most severe qualifications loads from each considered
launch vehicle. This static load determined for NORSAT-1 is 10.46g root-mean-square (Grms).
Applying a 5-sigma level of margin onto this load in order to account for various uncertainties in
the modeling, apply a level of safety margin, and instill a high level of confidence in the design,
results in a load of 52.3g. This load is applied as a global quasi-static acceleration in the FEM
model acting on the entire spacecraft, and is distinctly applied in each axis for six separate
loading scenarios. In order to reduce the number of simulation cases, the acceleration was
applied in three axes at once, resulting in just two separate cases to analyze with a 90.6g
magnitude load. Using this 5-sigma acceleration value of the random vibration represents a
conservative approach in analyzing the spacecraft structure, and avoids the necessity of a more
complex dynamic analysis.
4.3
4.3.1
Quasi-static stress results from simulations completed in June 2014, shortly after the CDR and
prior to part manufacturing, are compiled in Table 4 for each structural component. The relevant
requirement (#11 in Table 2) states that the spacecraft must have positive stress margins under
expected launch loads at 5-sigma. Stress margins are calculated as a measure of requirements
verification. Since the safety factor is incorporated in the applied load, a positive margin on the
yield stress for each component represents a satisfactory design. These margins of safety are
calculated as follows:
(4.1)
A margin of safety of 0% indicates that the design is just barely capable of withstanding the
applied loads (with safety factors) without failure, while a margin of safety of 100% represents
no loads being applied, and therefore would never result in failure. For example, for the
component +X Riser, the maximum applied load seen in the simulations is 161MPa, and the
50
failure load is represented by the yield stress of the material used (aluminum 6061-T6, 276MPa).
Using the above equation (4.1), a margin of safety of 42% can be calculated.
From the cases analyzed, all the stress margins are positive, therefore the design is acceptable. A
high level of safety margin is carried through in this analysis in order to instill confidence in the
structural design, since no physical structural tests will be performed on the spacecraft before
launch, other than an acceptance vibration test on the flight spacecraft at somewhat smaller load
levels.
The lowest stress margin is seen on the +Z tray near one of the –Y feet, a screen capture of this
result is shown below in Figure 33. As expected on both trays, relatively large stress
concentrations exist at all eight satellite feet, since these feet are the main load path to the rest of
the structure, and the only points of restraint in the XPOD-Duo. Small fillets were added to both
tray designs at the junction between the feet at the rest of the tray in efforts to minimize this
stress concentration.
Figure 33: Screen capture of –Z tray FEM stress results
51
4.3.2
The main requirement on the displacements seen under this worst-case load is to ensure that the
panel mounted solar cells would not see significant bending causing them to crack. Each solar
cell is limited to ~0.6mm displacement, which comes from a worst-case geometry based
calculation derived from the cracking radius of the cells (~2 meters). Allowable deflections on
each of the solar cell mounted panels are then determined geometrically based on this cracking
radius. In the analysis, all of the solar panels see a displacement significantly less than their
allowable deflections, with an average (plus 3-sigma) of ~0.37mm. Since all of the
displacements seen yield positive margins, they are acceptable. The displacements are also
monitored to ensure that no harmful contact is made between components themselves and with
the separation system. This is where the displacement requirements stem for the tray rails and
attachment brackets, the amount of displacement that would cause contact with the XPOD-Duo.
For all other components where no specific displacement requirement exists, a maximum value
of 1mm is used, which represents in many places, their distance to neighboring components, and
a conservative upper limit on their deflection. The displacement margins are calculated in a
similar fashion to the stress margins using equation (4.1).
Table 4: Stress and displacement analysis results summary
Component
Yield
Stress
(MPa)
Displacement
allowed
MAX
Displacement
(Absolute)
(mm)
Displacement
Margin (%)
MAX Stress
(MPa)
Stress
Margin
(%)
+Z Tray 276 1.00 0.17 83 223 19
+Z Tray Rail 276 0.41 0.17 59 223 19
-Z Tray 276 1.00 0.28 72 187 32
-Z Tray Rail 276 0.41 0.2 52 187 32
+Y Panel 276 2.82 0.21 93 192 30
-Y Panel 276 2.82 0.13 95 207 25
+X Panel 276 2.82 0.28 90 145 47
-X Panel 276 2.82 0.27 90 170 38
+Z Panel 276 2.28 0.2 91 142 49
-Z Panel 276 1.00 0.18 82 97 65
+X Riser 276 1.00 0.15 85 161 42
-X Riser 276 1.00 0.15 85 161 42
Separation Plate 276 1.00 0.19 81 85 69
Small Separation
Plate 276 1.00 0.14 86 54 80
52
Reaction wheel
Bracket 276 1.00 0.22 78 89 68
Large brackets 276 1.50 0.75 50 116 58
Half brackets 276 1.50 0.44 71 83 70
VHF Plate 276 1.00 0.21 79 52 81
Radio Cover 276 1.00 0.33 67 83 70
LP Electronics 276 1.00 0.2 80 48 83
LP Cassettes 276 1.00 0.17 83 71 74
CLARA 276 1.00 0.38 62 107 61
CLARA sun
sensor bracket 276 1.00 0.78 22 111 60
X Solar Cell
Coupons 276 1.00 0.25 75 83 70
Solar Panel Wings 325 2.82 1.2 57 157 52
Solar Panel Wing
Skin 591 2.82 1.2 57 359 39
Note that the maximum displacement values shown are absolute and at the system level,
therefore they don’t necessarily represent the amount a component is deformed. For example, the
displacement seen by the solar panel wings is largely due to the cantilevered method on which
they are mounted to the attachment brackets. The applied loads cause the brackets to deflect, and
the panels simply remain attached to the deflecting brackets. Through hand calculations based on
the relevant geometries, it can be found that the actual panel deflection, not due to the brackets,
is roughly 0.13mm.
4.3.3
Requirement #12 in Table 2 states that the spacecraft must have a first natural mode greater than
90Hz in order to comply with the worst-case launch vehicle requirements. Results from the
spacecraft modal analysis indicate that the first natural frequency (FNF) of NORSAT-1 is 144Hz
and is a local mode of the solar panel wing and attachment bracket oscillating back and forth as a
cantilevered mass. The result is shown below in Figure 34; it is above the minimum required
FNF of 90Hz with a 60% margin of safety and is therefore acceptable. The second wing shows
identical behavior at much the same frequency and numerous similar modes of the solar panel
wings and brackets are found around this frequency range. Note that the displacements shown in
the figure is not meaningful since this is a modal analysis with no specified loads.
53
Figure 34: Screen capture of simulation showing first mode of NORSAT-1 at 144Hz, and
subsequent mode frequency values
54
5
Honeycomb composite sandwich panels are commonly used in the aerospace industry due to
their numerous advantages over solid materials. Their high specific stiffness-to-weight ratio
allows them to be an extremely low mass solution for large structures where mass is of main
concern. In space structures, a common use for honeycomb panels is as large solar arrays – these
panels have a sole purpose of having numerous solar cells mounted on them for power
generation. The high stiffness of the panels allows them to be very simply supported, further
reducing their mass. Figure 35 below depicts their advantage over solid materials.
Figure 35: Relative stiffness and weight of sandwich panels compared to solid panels. Note
that the numbers shown are normalized to the solid material numbers [13]
In order to satisfy the power requirements required for simultaneous payload operations,
NORSAT-1 needs a total of six strings of eight solar cells (48) able to generate power during
operations of CLARA while pointing at the sun. Due to the spacecraft geometry proposed, the
largest face can accommodate at most two strings of eight solar cells, for a total of 16 solar cells
– less than half of what is required. If the satellite geometry were to grow in order to
accommodate the required cells, it would result in an inefficient mass structure, with large un-
used internal volume, and would no longer be compatible with the XPOD-Duo deployment
system. Because of this, it was evident that a dedicated solar panel would be required to meet the
requirements.
55
Due to a desire to minimize risks, fixed wing panels are favored over deployable panels for
NORSAT-1. Also, due to the size required to fit the 48 solar cells, it was decided to use a
honeycomb composite structure in an effort to not add significant mass to the spacecraft.
Initially, the proposed design for the large solar array matched closely to the ongoing SFL
project, NEMO-AM, which uses a large honeycomb sandwich panel as their main solar array and
is of similar size to NORSAT-1 (shown in Figure 36A). A second ongoing SFL project, NEMO-
HD, also uses honeycomb composite panels as various structural and solar panels (Figure 36B).
The panel design for NORSAT-1 is able to leverage the experience gained through these two
projects. This section outlines the design of the two large honeycomb composite solar panels
used on NORSAT-1.
Figure 36: NEMO-AM (A) and NEMO-HD (B) honeycomb panels
5.1
For NORSAT-1, there are very few specific requirements for the honeycomb panel design, other
than its main function; however, various programmatic drivers shaped the design approach. After
the initial proposal of a single large array that covers the entire spacecraft similar to NEMO-AM,
some preliminary thermal analysis concluded that it would be more beneficial to have a direct
line of sight to the spacecraft structure. From this, the panel was split into two separate solar
panel “wings”, which left the area of the satellite in between available for thermal control. The
NORSAT-1 “large solar array” refers to all of the solar cells located on the CLARA aperture side
(-Z side), which includes both front surfaces of the wings. Some additional general requirements
of these solar panel wings are explained below:
56
1) The large solar array shall be capable of successfully mounting 48 Azur 3G solar cells.
o This is the only size-driving requirement on the panels, however they are also
kept as small as possible in order to minimize the overall bus volume.
2) The large solar array shall meet the same structural and modal requirements as set for the
entire spacecraft (positive stress margin, first natural frequency greater than 90Hz).
o This ensures it will meet the system level requirements, however, because the
panels are not entirely independent, their design is largely interrelated with the
entire structure and will be analyzed together as well.
3) The large solar array shall be able to withstand temperatures ranging from -100°C to
+80°C
o Predicted temperature range from the thermal model. This range is large because
the panels are thermally isolated from the spacecraft bus.
4) The large solar array should be made of common components/materials that do not
require a long lead time, and should be composed of similar materials to that used by
NEMO-AM and NEMO-HD
o Schedule is one of the main concerns for the project. Because the panel is a fairly
simple design, the details should not be over complicated, resulting in a long lead-
time. Using heritage components from previous projects can also reduce some of
the risk associated with the foreign materials.
5) The two solar array wings should be identical.
o This is done in order to save money during procurement of the panels, and allows
for a modular design.
6) The large solar array should be designed such that it can be integrated onto the satellite
last.
o This is done to facilitate the integration process of the satellite, and to minimize
the handling of the spacecraft while these expensive/delicate panels are installed.
A large goal in the panel design was to not add significant cost and time to the project; the panels
were kept as simple as possible in order to achieve this. For example, additional components,
such as antennas and sun sensors, were excluded from being mounted on the panels, for further
complication of their design. As was shown in Chapter 4, the mass of these panels directly affect
57
the first natural frequency of the spacecraft, therefore, extra effort is placed in keeping the mass
of each panel as low as possible and accurately predicted.
5.2
Derived from the above requirements, and fitted to the relatively mature spacecraft structural
design, a design concept for the honeycomb solar array was realized. Two identical “wings”
make up the entire solar array, each supported by five mounting inserts, attached to the satellite
via three aluminum attachment brackets. The size of each wing is 20cm x 50cm, so as to each fit
three full strings of eight solar cells, to make up the entire array of 48 cells; additional area is left
for the wiring of the cells, as well as ground support equipment (GSE) holes for various panel
protectors and handling measures. Although the 50cm dimension is larger than the longest
dimensions of the satellite bus, the panels are mounted slightly off-center in order to allow the
spacecraft to rest on its –Y face, and facilitate the installation of the panels at a late stage. Also,
to assist in the late installation, each panel is fitted with a dedicated connector to carry the wiring
of each panel’s solar cells, and a single temperature sensor, into the spacecraft avionics; this
connects to a filtered skin connector on the –Z tray into the avionics bay of the spacecraft. In
order to keep the panels as small as possible, the mounting inserts used are blind threaded holes;
this allows the front side of the panel to be fully available for the placement of the solar cells. A
depiction of this proposed concept is shown in Figure 37. A single string of eight solar cells is
also placed centered on the back of each panel, to act as the –Z face solar array. An extra insert
was included in order to keep the panel design and solar cell placement/wiring identical on each
panel.
Figure 37: NORSAT-1 solar array concept design
58
The attachment brackets used stem from the attachment bracket design used on the NEMO-AM
mission, mounting to the spacecraft and the panels in a similar fashion. However, due to the
separation of the large panel into two small panels, there exists a cantilevered mass of each of the
panels on the extended portion of each of the brackets, which will result in a low frequency mode
of the satellite. In light of this, the brackets are reinforced in the direction of bending by using an
I-beam concept to keep the mass as low as possible; this is shown in Figure 38. The appropriate
sizing of these brackets is designed through an iterative process during the finite element analysis
of the entire spacecraft in order to maximize the first natural mode; this was detailed in Chapter
4.
Figure 38: Wing attachment bracket design
The remaining honeycomb panel specifics, such as core/skin thickness and other relevant
honeycomb parameters, are designed based on failure criteria of the sandwich panel and insert
design.
5.3
Due to the complexity of sandwich panels, several failure modes exist as opposed to general
isotropic materials. Separate failure modes exist for the sandwich panel itself, and for the inserts
that are added for fixation purposes.
5.3.1
Sandwich panels have many failure modes as described in various literature, including in [13]
and [14], which were referenced heavily during this study; some of these are briefly described
below and in Figure 39.
59
a) Face yielding/fracture - Tensile or compressive failure of the face sheets
b) Core Shear – Core material shear failure
c) Face wrinkling - Local face sheet buckling inwards due to core compression failure
d) Delamination – Local face sheet buckling outwards due to delamination
e) General buckling - Panel buckling due to in‐plane compressive loads
f) Shear crimping - When buckling results in additional, localized core shear failure
g) Face dimpling - Multiple inter‐cell buckling due to large core material cell sizes
h) Core Indentation – Failure of the core due to localized pressure
Figure 39: Sandwich panel failure modes [15]
All of these failure modes were investigated for the NORSAT-1 honeycomb panel wings,
however, due to the expected loading on the panels many of these modes are not applicable. For
example, failure modes c), d), e), f) and g) only occur when large compressive loads are applied
to the panels. Due to the proposed method of mounting, the NORSAT-1 honeycomb panels will
be simply supported by some fixation inserts, and will carry no loads other than their own weight
under a directional acceleration, therefore, will see very little to no compressive loads. The
remaining three failure modes, face/core yielding/facture, core shear, and core indentation, were
analyzed as part of the panel design. These failure modes, however, are not expected to be design
drivers. Due to the mounting method of the panels and the expected loads, insert failure is likely
to dominate. The equations for determining the relevant stress via these three failure modes are
displayed below for a simply supported honeycomb panel.
60
Facing Stress Core Shear Stress Local Compression
(5.1) (5.2) (5.3)
Where:
M = Maximum bending moment
h = Distance between facing skin centers
tf = Thickness of facing skin
b = Beam width
F = Maximum shear force
P = Applied load
A = Area of applied load
The deflection of the panel is also of interest due to having sensitive glass covered solar cells
directly mounted, which can crack if a large enough deflection is seen. The deflection of a
simply supported honeycomb panel “beam” can be expressed as:
(
) (
) (5.4)
Where:
kb and ks = Deflection coefficients from [13]
D = Panel bending stiffness
S = Panel shear stiffness
l = Panel free length between supports
The above equations are used to validate the NORSAT-1 proposed panel make-up, and the
relevant results and margins of safety are displayed in Table 5 of Section 5.4. The boundary
conditions of the panel are simplified to approximate a simply supported beam panel with an
applied load. The fifth center mounting point of the panels is ignored, and the two mounting
points on either end of the panel are approximated as the simple supports, to approximate a
scenario as shown in Figure 40.
61
Figure 40: Simplified loading scenario for sandwich panel failure mode calculations [14]
5.3.2
A common failure occurrence in sandwich structures is through insert failure. Inserts are the
main method for adding fixation points to sandwich panels. This is often done by a post-cold
process, whereby an insert is implanted into an already laid-up sandwich panel and fixed using a
two-part epoxy resin system, or “potting compound”. Because of this post-process
manufacturing, they often tend to be a weak point of failure, depending on how they are being
used, because of the discontinuity placed in the sandwich panel. Although significant work has
been done to predict the behavior of these insert systems, they continue to be a somewhat fickle
point of failure because of their sensitivity to the small details in the manufacturing process.
Insert pullout tests are sometimes done in order to characterize the insert installation method,
however these are often forgone for small projects due to their expensive nature, and panels are
instead often compared to relevant similar test data. Generally, predicted insert failure through
analytical methods prove to be a conservative approach when compared to actual test data,
barring any manufacturing defects [16].
There are four main types of insert loading scenarios, which can each lead to various failure
modes of the face sheets, potting compound, honeycomb core, or insert itself, due to the added
discontinuity in the panel. These are displayed in Figure 41 below, and are each discussed in this
sub-section. The Shur-Lok and ESA Insert Design handbook ([17] and [18] respectively) were
referenced heavily during this study.
62
Figure 41: Insert loading scenarios [18]
Loading scenario a) involves an out-of-plane tensile or compressive load. A tensional load leads
to a failure mode commonly known as pullout, which depends greatly on the core thickness of
the honeycomb and has little effect from the face sheets. An expression provided in [18] for
determining the critical pull out force ( ) of a partially potted (blind hole) insert in a
sandwich panel where the core thickness is much larger than the skin thickness is shown below.
A diagram of the scenario is also shown in Figure 42.
( )
(5.5)
Where Pcrit is the static load carrying capability for a fully potted insert under the same
conditions, evaluated by:
(5.6)
And where:
bp = Effective potting radius
hp = Potting height
c = Core thickness
τccrit = Core shear strength
σccrit = Core tensile strength
d = Distance between the face sheet middle surfaces
63
Figure 42: Partially potted insert under tensile load [18]
An effect of the insert’s distance to the edge of the panel also accounts for an appreciable
reduction in strength of the insert. This factor is incorporated into relevant calculations through
means of amplification on the applied load, as detailed in the Edge Influence section of [18].
For a partially potted insert under out-of-plane compressive load, the critical failure load is
almost always larger than the critical failure load under tensile load, due to the added strength of
the bottom face sheet and the increased compressive over tensile strength of the core, therefore
its calculation is not considered for the NORSAT-1 panels.
Loading scenario b) involves in-plane shear loads on the insert, which can result in various
failure modes of the face sheets and core material. For a sandwich panel with CFRP face sheets
and aluminum honeycomb core, a permissible shear load ( can be determined based on the
combined properties of the core and face sheets, this formula is shown below for potting radius’
(bp) less than 11mm [18]. In addition, four independent failure modes of the CFRP face sheets
can occur: Tensile, Shear-out, Dimpling, and Bearing failure; these are shown in Figure 43.
(5.7)
Where:
bp = Effective potting radius
τWcrit = Core shear strength in (weaker) W direction
σfy = Yield strength of face sheets
f = Thickness of upper face sheet
64
Figure 43: Face sheet in-plane insert failure modes [18]
The detailed predicting formulae of the above face sheet failure modes are comprehensively
explained in the Insert Design Handbook [18]. Due to the relatively high strength of CFRP, these
failure modes tend to be quite insignificant, but depend heavily on the chosen thickness of the
face sheets. Another large influence is distance from the insert to the edge of the panel – this has
a substantial influence particularly on the shear-out failure mode.
Loading scenario c) involves an in-plane torsional load to an insert, which could lead to shear
failure between the potting adhesive and honeycomb core. This is greatly minimized through the
use of multiple inserts, and thus is not a concern for the NORSAT-1 panels. The only torsional
load that will be seen would be the screw torque, which will be substantially less than the failure
torque.
Loading scenario d) involves an out-of-plane bending moment placed on the insert. Inserts are
typically not advisable to be placed under such loading conditions because of their poor
performance under such conditions. Much like reducing the torsional load, this is commonly
mitigated through the use of coupled inserts, which can convert the bending moment to a
tensional and compressive load. It is further mitigated through the use of a clamping surface (or
washer) that is larger than the insert diameter, capable of spreading the load across the face sheet.
In light of these simple mitigation methods for failure through torsional and bending loads, these
insert failure modes were not considered for the NORSAT-1 panels.
65
5.4
Given the proposed design concept and the relevant failure modes, the specifics of the
honeycomb panel for NORSAT-1 can be realized. Figure 44 shows a breakdown of the resulting
sandwich panel design.
Figure 44: NORSAT-1 solar panel wing design
The panel specifics were derived through trial and error, in an effort to minimize mass, obtain
positive safety margins, and make use of standard sizes and parts (i.e. stock inserts, common
CFRP ply thicknesses and orientations, standard honeycomb core details). Aerospace grade
honeycomb core material is used, which is commonly either Al 5052 or Al 5056, and
incorporates vented honeycomb cores that make them appropriate for use in vacuum. Carbon
Fiber Reinforced Plastic (CFRP) skins are used due to their low-weight, and minimum distortion
in extreme temperatures – this matches closely with the solar cells, which have a glass cover
component with very low thermal expansion. In minimizing the difference in thermal expansion
of the face sheet and the solar cells, it also minimizes the potential stresses that could occur due
to mismatch in thermal expansions and cause solar cell cracking. Based on these panel details, an
estimated mass of each panel, prior to solar cell laydown, is approximately 300 grams.
In arriving at these panel details, the relevant failure mode margins are shown below in Table 5.
In all cases, the applied gravity load is matched to the load used in the finite element analysis of
the spacecraft (Chapter 4), which covers all expected shock values, taken as 52.3g’s. The load is
applied in the worst-case direction for each failure mode, and is distributed among four of the
inserts, neglecting the middle insert, and assumes each insert has the worst-case edge influence;
Honeycomb Panel Wing Design Details
Honeycomb core material,
cell size, density, foil
thickness
Al 5052, 4.0mm, 3.8pcf,
0.03mm
Honeycomb core
thickness
9.5mm
Skin material, density, #
of plies, ply orientation
CFRP, 1446 kg/m3, 2 ply,
0/90
Skin thickness 0.25mm
Insert material, Type,
diameter, Height
Al 2024-T4, Fully Potted,
11mm, >6mm
66
combined loading was not considered. These margins of safety are calculated similarly to as
described in Section 4.3.1 using equation (4.1); since the safety factor is incorporated into the
applied load, a positive margin represents a satisfactory design.
Table 5: NORSAT-1 honeycomb panel design failure modes summary
Panel
Failure mode
Facing Stress
Core Shear Stress
Local Compression
Deflection
Margin of Safety
96% 93% 99% 76%
Insert
Failure mode
Critical Shear
Tension Shear-out Dimpling Bearing Pullout
Margin of Safety
87% 98% 64% 30% 94% 83%
As seen, all of the calculated margins of safety are positive, and therefore the panel design is
satisfactory. The most critical and relevant failure modes proved to be out-of-plane insert pullout
and in-plane insert shear-out. The pullout failure is most heavily dependent on the core thickness
and insert size, where having a thicker core, and larger diameter insert would result in a higher
resistance to failure. However, the shear-out failure is most dependent on the edge influence
caused by the location of the insert relative to the panel edge, where a larger diameter insert
would only make this distance smaller, and more prone to failure. Therefore these variables had
to be balanced to ensure acceptable positive margins in all affected failure modes.
All other modes of failure proved to be fairly irrelevant due to the low expected loads on the
panel. The deflection of the honeycomb panel was found to be as low as 0.9mm, which is well
below what would be required to cause the panel mounted solar cells to reach their cracking
radius of two meters. Although the insert dimpling failure mode is seen to be relatively low, this
failure mode is difficult to predict analytically due to its dependence on the specific panel
geometry, type of loading, and boundary conditions and is usually determined through test. As
seen, the panel was not optimized down to minimum margins; this is due to the relevant
uncertainties in the panel manufacturing that cannot be anticipated, thus higher margins are kept
in order to reduce risk and instill additional safety into the design. In addition to this, very small
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mass and cost savings are gained through reducing the panel size further, therefore a slightly
over-designed panel has little negative impact on the project.
Also exists is an abundance of data regarding the tensile and compressive strength of post-cold
bonded inserts in [19]. The data is presented in graphical form, where a sandwich panel of
particular make-up and insert size is considered, and shows how the minimum and average insert
strength values vary as a function of the core height. The data is produced by means of an
analytical method, though has been compared and validated through empirical testing [20].
According to [19], the minimum and average tensile strengths displayed are applicable to
honeycomb panel design without further investigation, provided that the core properties match,
the manufacturing of the inserts were performed in accordance with their specified quality
assurance, and the load is carried by multiple inserts. Although not certain that the quality
assurances of the NORSAT-1 panel will match, and the different facing material used, given that
the main parameter of the pull out strength is the core properties, these values can be regarded as
ballpark values. The graphs for tensional and compressive load on a similar honeycomb core and
similar sized insert are shown in Appendix A. As seen, from these graphs, the minimum pull out
strength capability is above 500N in both cases, and the typical values are near 1000N. These
values relate fairly close to the predicted worst-case pullout load for the NORSAT-1 panel
inserts, which was found to be approximately 450N per insert.
5.4.1
Because the NORSAT-1 panel isn’t very sensitive to small changes in the honeycomb make-up
details, procurement of the panels involved compromising on certain details in favor of
items/materials that were in stock, which would result in quicker panel procurement. For
example, the panel manufactures suggested switching the insert material from Al 7075-T73 to
AL 2024-T4 due to them finding the latter material in stock. In each case, the relevant change
did not significantly affect the failure modes of the panel, and would only make the panel more
robust. Some of these changes would result in additional mass over the original panel design or
modifications to the attachment brackets; however it was deemed that the schedule would take
priority. The NORSAT-1 honeycomb panels are currently being manufactured and discussions
regarding various material/design compromises are on-going with the manufacturer.
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6
In an effort to reduce the overall launch volume of NORSAT-1, many of the required large
protruding components were made deployable. By reducing the overall volume via deployable
components, NORSAT-1 becomes easier to accommodate as a secondary payload aboard a third-
party launch provider. On NORSAT-1, two very high frequency (VHF) receiving antennas, and
four Langmuir probe booms have been made deployable, for a total of six deployed components.
6.1
The Space Flight Laboratory (SFL) and the Norwegian Space Centre (NSC) have collaborated
on a number of Automatic Identification System (AIS) satellites. Numerous previous satellites
on-orbit of the GNB size, including AISSat-1 and AISSat-2, use a single pre-deployed VHF
antenna as shown in Figure 45. The pre-deployed method being favored because of the reduction
of risk over deployable components (i.e. a deployable component always has a single point of
failure – if it does not deploy properly). The size of the GNB satellite bus is less than half the
volume of the NEMO bus type, causing the added volume of the antenna to be less influential on
the restrictive nature of the launch volume.
Figure 45: AISSat-2 pre-deployed VHF antenna approximate dimensions
In the case of NORSAT-1, however, two orthogonal antennas are required, rendering the volume
consumed by pre-deployed antennas to be significantly larger. The added volume by two
orthogonal antennas of each ~50cm length would restrict the number of potential launch
opportunities, and would place additional constraints on the launch vehicle in order to
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accommodate the satellite. Because of this, deployable antennas were favored over pre-deployed
for the two VHF antennas on NORSAT-1.
6.1.1
Tape springs are a common solution to deployable components in space applications, due to their
lightweight, energy storage, and stiff characteristics. An on-going GNB sized satellite project to
be launched in the near future was already at a mature design stage at the time of decision, and
uses a tape spring system as a deployable VHF antenna (EV-9). Because of SFL’s heritage and
experience in their design and use as VHF receiving antennas, tape springs are used as the
deployable VHF receiving antennas for NORSAT-1 as well. No other deployable antenna
systems were considered, because of the believed favorable simplicity of the tape spring
approach.
Only three main requirements exist for the antenna design:
1) The antenna shall be a simple 50-Ohm quarter-wavelength (~46 cm) monopole with no
specific beam width.
2) The two antennas shall be mounted orthogonal to each other (Requirement #9 in Table
2).
3) The deployment shall be able to be performed and tested in a 1g environment
(Requirement #14 in Table 2).
The first requirement limits the antennas to materials that can be used as antennas (conductive
materials), and also sets a length to achieve the necessary frequency (162 MHz), while the
second requirement specifies their orientation. In addition to these, the antennas should also be
easily stow-able, and able to be installed last, in order to simplify ground handling and testing.
An additional desire for the deployment design is to have the antennas deploy passively, not
requiring additional equipment such as burn wires or on-orbit commanded deployment systems –
this can significantly simplify the design and reduce the risk of failure.
6.1.2
Research was conducted to better understand tape springs and use them as intended, which
ultimately shaped many of the design decisions. These activities are listed and explained below.
Most of these activities were explored using Commercial Off The Shelf (COTS) carbon steel
tape measure tape springs.
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Width
Typical tape measure tape springs come in two widths: one-inch and half-inch. The difference in
width have little effect on the antenna performance, however the larger width tape spring has a
larger mass, causing gravity to have a larger effect on it during testing on the ground.
Machining
Machining a tape spring can be a difficult task, due to its extremely thin and curved cross
section. For cutting the tape spring to size, it was found that a sharp pair of shears would do the
best job, while a duller pair may cause the corners of the cut to become slightly deformed, and
would affect the continuity of the tape spring. When making holes, a drill press was found to be
the preferred method, over other methods such as hole punching. Best results were achieved
when drilling into the concave side of the tape spring, and ensuring a hard piece of material was
below the tape spring. Clamping of the tape spring to the hard material as close as possible to the
location of the drilled hole was also necessary to ensure the tape spring would not get caught in
the drill bit blades. Punching the hole-location with a center punch prior to drilling also helps a
great deal to ensure a smooth drilling process. An improperly drilled hole in the tape spring can
easily distort the continuity, and will cause the tape spring to perform undesirably.
Critical Bend Radius
Since the method of deploying tape springs often involves bending them in some fashion, and
allowing them to store the energy needed to deploy, critical bending limits were explored.
Through testing on COTS half-inch tape measure tape springs, the following bending limits were
found. Bending was performed with the concave side of the tape spring facing inwards to the
bend, as shown in Figure 46.
Figure 46: Tape spring bend radius
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< 4mm radius = plastic deformation.
< 6mm radius = No visible deformation or decrease in performance, however the tape spring has
a preferred kink spot when bent in that area.
> 6mm radius = No change in performance.
Materials
All COTS tape measures are made of carbon steel, and often have a yellow plastic coating on
them for annotations - these materials aren’t ideal for space applications. The plastic coating can
easily be removed using a paint remover and/or through abrasion. The carbon steel material is of
slight concern, due to its magnetic properties. The dipole created can affect the readings from the
on-board magnetometer used for attitude sensing. It is because of this concern, that another
ongoing SFL project, CanX-7, has opted to create a custom tape spring in-house, out of non-
magnetic Copper Beryllium (CuBe), shown below in Figure 47. Their use for the tape springs,
however, is slightly different than for NORSAT-1. On CanX-7, these tape springs are used for
deployment of a large drag-sail, and are stowed in a very small containment unit, having the tape
spring wrapped around itself several times. Once deployed, the booms are more than double the
length of the planned NORSAT-1 VHF antennas, and thus, if steel tape springs were used, they
would cause a larger dipole to be seen which affects the performance of the limited attitude
determination hardware on board.
Figure 47: Tape spring materials explored
Though the same curvature radius was created, the CuBe tape springs were found to have
approximately 20% less stiffness than their steel counterparts. This causes the antennas to be
more influence by gravity during ground testing. Additional leftover CuBe tape spring material
was only available in the one-inch width size, and it was not desired to go through the effort of
time/money to manufacture thinner ones.
Stowage
A similar stowage technique to that used on EV-9 (using a tooling ball) was desired, however, in
NORSAT-1’s case, there were two antennas to hold down rather than just the one. The
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difficulties of holding down multiple tape springs were explored in order to potentially hold the
antennas together using one single technique. Considerations were to ensure a relatively stiff free
length of antenna, to avoid accidental contact with the spacecraft, and rattling between the two
antennas during vibrations. Four options existed when wrapping the antennas together,
depending on the direction of each antenna’s curved face – these are depicted below in Figure
48.
Figure 48: Hold down configurations for two tape spring antennas
It was found that scenario 1, with each tape spring nestled together, with the concave face facing
downward, was the best solution. By having both antennas nestled in one-another, there is little
chance for rattling/slippage between them during vibrations, and both antennas can act as one. It
also allows the antennas to bend in their preferred direction, curved face down, around the
satellite, minimizing chances of plastic deformation when bending. Having the curved face
downward also allows the antennas to be flattened out at the hold down point, causing the
stiffness of the free lengths to seemingly increase.
Relative Placement
It was desired to have both antennas be held down at a common point to allow for the simplicity
of a single deployment mechanism; this may require the antennas to be placed quite close to one
another, and could cause some negative antenna coupling effects. In order to assess the
performance of the antennas at various locations on the satellite and distances from each other,
Clement Ma, communications engineer at SFL, performed a number of simulations of the
orthogonal antennas at various placements on the spacecraft. It was concluded that a favorable
configuration has the antennas mounted on a common face (+Z panel), with a distance of 100mm
between them, and angled 45° from the spacecraft.
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Mounting
As learned from the EV-9 project, when mounting the tape spring to a flat surface, the tape
spring will flatten out at the mounted point, and cause it to lose stiffness and not perform as
expected. To solve this problem, the tape spring can be mounted to a curved surface, matching
the tape spring’s curvature; this causes no loss in continuity. However, it was later found that
when tightening the mounting screws a top the tape spring onto the curved surface, if the
direction of the mounting screws do not match the curved surface, the screw head will not make
full contact, and may deform the tape spring under the screw head. This would affect the angle of
the mounted antenna. Having the mounting screws at an angle matching the curve would solve
this problem, but would introduce more complicated machining of the antenna mounting piece.
6.1.3
The design of the deployable AIS antennas is derived from a previous SFL designed AIS antenna
used on the GNB. Tape spring antennas are used, are wrapped around the satellite and held down
at the top +Y face by a spherical tooling ball attached directly to the XPOD. Upon release, the
antennas deploy themselves by using their own stored energy. This technique can be seen in
Figure 50, which shows the satellite installed in the vertically mounted XPOD-Duo.
In light of some of the conclusions from the research done using the tape springs, each tape
spring is mounted differently, such that when wrapped together their curved sides are concentric
and can rest together. The concave side of the tape spring is faced towards the satellite to avoid
accidental contact to the spacecraft (this is the stiffer direction), to allow the tape springs to wrap
in their favorable direction, and be stiffer at the hold down point; as explained in the research
section above. Minimum bending radii are limited to 7mm to avoid any damage to the antennas
during stowage.
COTS carbon steel tape measure tape springs are used, and are machined in-house by the author.
It was deemed that the relatively small dipole introduced by the magnetic steel is of not large
concern, and it was preferred to have the stiffer, lower mass half inch steel tape spring to
facilitate ground testing. This is also common to the material of the antenna used on EV-9.
SubMiniature Version A (SMA) connectors are used as the input for each antenna, and are
soldered to each VHF mount through a seal feed through component. Once soldered, the sub-
assembly of the VHF mounts and the mounting plate remain together. Each VHF mount is
coated in gold to increase its conductive ability. The antennas are mounted 100mm from each
74
Stowed AIS
Antennas
other, in accordance with the antenna simulations, and at 45° to the satellite, using uniquely
designed VHF mounts, with either a concave or convex mounting surface for the antenna. The
design of the specific components of the antenna base is largely based on the one used on EV-9,
however the components were miniaturized and simplified where possible. The antenna base
design for NORSAT-1 is shown below in Figure 49.
Figure 49: NORSAT-1 antenna base exploded view
Originally, when using the old vertically mounted design of the XPOD-Duo deployment system,
the antennas would be held down at the top +Y face by a spherical tooling ball attached directly
to the XPOD door. A separate small bracket, as seen in Figure 50, would attach to the XPOD
door via two screws, and this bracket would house the tooling ball. This method allowed for
minimal modifications to be made to the XPOD (just requiring two tapped holes to be drilled),
and matches closely to the deployment technique for EV-9. The antennas would then deploy
upon release of the XPOD door, and would deploy outward, away from the launch vehicle.
Figure 50: NORSAT-1 in the vertically mounted XPOD-Duo
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However, when the XPOD-Duo design was adjusted to mount horizontally, this deployment
technique had to be reevaluated. Due to the large solar array wings, the orientation of NORSAT-
1 in the horizontally mounted XPOD-Duo is restricted, which renders the VHF antennas slated to
be deployed downward, towards the launch vehicle, assuring that some sort of contact would
likely be made (depending on the geometry of the mounting surface). In talking with the launch
providers, having the antennas potentially contact and scrape the launch vehicle mounting
surface would not be a large concern, and would also not be a large concern for damaging the
antennas, due to their extremely low mass and force of deployment.
Two possible options were explored: leaving the orientation as is, with the hold down point on
the XPOD door, or reorienting the satellite 180° in the XPOD and having the hold down point on
the XPOD pusher-plate rather than the door. With the former, the antennas would be travelling in
the opposite direction of the satellite, and would therefore have a chance of catching on
something, and potentially obstructing the deployment or damaging the antenna. With the latter,
the antennas would deploy in the same direction as the satellite, and would tend to simply slide
on the bottom surface in a favorable direction. In both cases, there is still a large chance that the
antennas will make contact with the launch vehicle. Because it was deemed acceptable to have
the antennas contact the launch vehicle, the latter option was favored, in order to ensure a smooth
deployment.
A small bracket was designed with Michael Ligori, one of the XPOD-Duo designers, which
would house the tooling ball, and attach directly to the pusher-plate via two attachment screws.
The bracket is mounted on the exterior of the plate, and is accessible on one of the open faces of
the XPOD – this way, deployment stowage can easily be inspected, and the bracket can also be a
late addition to the assembly. The bracket proposed is shown below in Figure 51. The tooling
ball fits into a carefully sized hole on the antenna hitch of the +Y panel on the satellite, while
securing both antennas through similar sized holes. The use of a spherical tooling ball is ideal to
ensure no obstructions can be made whilst releasing, also, Antenna 2 is fitted with a slightly
larger hole than Antenna 1, in order to ensure contact with the tooling ball and avoid slippage.
The antenna hitch and guides are made of Delrin plastic so as to minimize potential damage
during vibrations due to friction.
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Figure 51: VHF antenna stowage system
A simple stowage technique is also applied to the antennas, whereby a standoff can be threaded
through both of the antennas into the antenna hitch, locking the antennas in place until
integration with the XPOD. Once loaded in the XPOD, the standoff can be removed by hand
from the opening, and the antennas are then held solely by the pusher-plate. A large, noticeable,
“REMOVE BEFORE FLIGHT” tag will be added to this standoff to ensure that it is not
forgotten to be removed before launch. An additional antenna guide is also placed on the +Z
panel, in an effort to reduce the free length of the antennas while stowed. This guide is taller than
the one shown in Figure 51, in order to ensure that the antennas physically bend over it,
increasing their stiffness. Figure 52 shows the antennas stowed on the spare structure of the
spacecraft, and the relevant antennas guides.
Figure 52: NORSAT-1 side view of stowed AIS antennas on spare structure
77
Thermal considerations are still underway for these VHF antennas, however, the current thermal
solution is to strip the yellow paint off the antenna, and apply Kapton tape directly to the steel
faces – this is to keep the antennas from becoming too hot in orbit due to their extremely small
thermal mass, and also prevent the antennas from rusting on the ground.
The ideal length of the quarter-wavelength antennas can be derived by using the wavelength
formula, where the wavelength = the speed of light ÷ the frequency. Using the speed of light in a
vacuum, and the frequency of interest (162Hz), 462.6mm is obtained. The antennas for
NORSAT-1 are cut to a length slightly longer than this (482.6mm), in order to allow for tuning
to be done on the flight spacecraft before launch, or during VHF antennas pattern testing – this
way they can be easily cut shorter at a late stage if necessary.
Deployment tests were performed in order to better understand the behavior of the tape spring
antennas in the proposed deployment method, and confirm the assumptions being made. The
conclusions from these tests are explained in Section 8.6. A depiction of the expected behavior
of the antennas is shown below in Figure 53, as the spacecraft is being ejected from the XPOD-
Duo. Having the constant pushing force of the pusher plate upon deployment, the antennas are
expected to remain stowed until the deployment spring of the pusher plate has reached its free
length (which is about three-quarters of the length of the XPOD), after which the antennas will
be able to slip out and deploy.
Figure 53: Expected deployment volume of VHF antennas
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6.2
The Langmuir probe deployable mechanism was designed and tested by the payload provider,
the University of Oslo. Brief details of the design are mentioned in Section 2.3.2. Two separate
cassettes house two Langmuir probes each, each mounted on a long boom. The booms are
deployed using a Shape Memory Alloy commercial pin puller on orbit via a ground command,
and are each locked in place with a locking pin once fully deployed. A simple position sensor is
included to confirm a successful deployment. Figure 54 below is a depiction of the probes
deploying from their cassette.
Figure 54: Langmuir Probe cassette deployment
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7
Ground support equipment (GSE) includes anything that must be used in order to facilitate the
development of the satellite from beginning to finish. This can include anything from software
for testing, to shipping containers used to transport the satellite to the launch site. In accordance
with the micro-space philosophy, one of the main driving requirements for all GSE designs is
that they be kept simple and low cost. The author was mainly involved with the mechanical GSE
(MGSE) needed to support assembly, handling, and some of the mechanically related testing –
this MGSE developed for NORSAT-1 is detailed in this section.
7.1
In order to reduce the risk of damaging components during assembly and integration of the
satellite, various MGSE is designed to ease the assembly process. Some requirements for these
designs are:
1) The Assembly and Handling GSE should be multi-use wherever possible, in order to
minimize cost and complexity
2) The Assembly and Handling GSE should interfere with the final assembly or be distinctly
labeled, in order to avoid forgetting to be removed
3) The Assembly and Handling GSE should use GSE-specific holes wherever possible, in
order to reduce the number of cycles on flight screw holes
4) The Assembly and Handling GSE should be lightweight, in order to not add significant
weight during manipulation of the satellite
5) The Assembly and Handling GSE should be easy to install/remove, and be low risk of
damaging sensitive components
6) The Assembly and Handling GSE should not load the honeycomb panels
One example of this is a set of four Delrin machined tray “legs” that were designed to allow
clearance for mounted components in multiple orientations of the +Z tray during assembly, as
well as to ease the process of connecting the +Z tray to the –Z tray prior to any panel integration
– this is depicted below in Figure 55. Additional GSE was designed to facilitate the installation
80
of the +Z panel while complying to the order of assembly, as well as support stands for all sub-
assemblies (such as panels), during sub-assembly integration.
Figure 55: MSGE assembly tray “legs” in various orientations of use (black)
Due to the many external protrusions of NORSAT-1, finding a comfortable resting orientation
that eases assembly and disassembly proves difficult. In order to have the large solar array wings
away from harm, as well as both the deployed AIS antennas and deployed Langmuir Probes free
from obstruction, the natural resting position has the +Z face towards the ceiling, and the –Z face
with the solar array wings resting on the ground or work bench. This position can be easily
achieved prior to integrating the protruding wings and CLARA payload by resting the satellite on
the solar array wing attachment brackets as seen in Figure 55 above. Once these components are
integrated however (near the end of the assembly procedure), we can no longer use this resting
technique, and additional GSE is needed.
A support stand was designed to solve this problem and is shown below in Figure 56. Using two
end-platforms and a pair of cross braces, it delivers a sturdy stand to rest the satellite on
throughout the assembly, integration, and testing. The satellite rests directly on the end platforms
using the solar array wing attachment brackets, and is secured to the platform using three M4
sized thumbscrews on each side. The support stand raises the satellite ~12cm of the ground
(work surface), and is also used for deployment testing, and integration into the XPOD-Duo
deployment system (detailed in section 7.5). Copper straps are added to all the Delrin handling
81
GSE to ensure grounding of the satellite during integration; an example of the copper strap is
seen in Figure 61 and Figure 66 on the support stand and assembly legs.
Figure 56: NORSAT-1 GSE support stand
Due to the awkward geometry and relatively large mass of NORSAT-1, handling of the satellite
during and after assembly can be a difficult task and must be done with the utmost care to avoid
potential mishaps and damage to interior and exterior components. A set of GSE handles was
designed to facilitate the transport and manipulation of the satellite during/after assembly and is
shown below in Figure 57. The overall strength requirement for this handle system is that it be
capable of supporting the entire mass of the satellite, and any protective or supporting GSE under
a load factor of 5g.
The handle design is adapted from a design used for the NEMO-AM project, and employs a
thick aluminum support bar that directly attaches to the wing attachment brackets into GSE-
specific threaded holes. This provides a reliable pick-up point on either side of the satellite on
which the remaining handle assembly is then mounted, consisting of a half-inch aluminum bar
and three aluminum brackets. This handle assembly can be quickly installed and removed from
the satellite when needed using the three M4 thumbscrews, and provides an unobtrusive two-
person system for maneuvering and handling the satellite. The support bar is also used to provide
pick up points for thermal vacuum testing and craning via eyebolts. An alternate handling system
is also incorporated directly on the support stand, using commercial offset handles, to provide
additional pick-up points that would be favored in certain scenarios – these can be seen in Figure
61, and were mainly added to facilitate loading the spacecraft into the XPOD- Duo, and
deployment testing.
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Figure 57: NORSAT-1 GSE handle assembly
7.2
Some of the most sensitive and expensive components on a typical microsatellite are mounted
directly on the exterior surfaces – these are mainly the solar cells. These cells have a glass top
layer that can be easily shattered with a sharp impact. Also on the exterior are several delicate
thermal tapes that could be damaged through constant unnecessary contact. Because of these
reasons, among many others, such as keeping the satellite as clean as possible, a protective
enclosure is designed to house the satellite after assembly. With access to the testing port of the
satellite, much of the pre-flight testing can still be done on the satellite whist the protective
enclosure is installed, minimizing risks of physical damage to the satellite up until integration
into the deployment system and onto the launch vehicle. Some requirements for the protective
enclosure design include:
1) The protective enclosure shall be clear, in order to allow visual inspection with the
enclosure installed.
2) The protective enclosure should need minimal tools to install/disassemble.
3) The protective enclosure shall provide a barrier over all solar cells and delicate exterior
components.
4) The protective enclosure shall allow for access to the testing port of the satellite.
5) The protective enclosure shall be composed of static dissipative materials.
The protective enclosure design for NORSAT-1 is based on the protective enclosures previously
developed for SFL’s GNB class of satellites. It consists of multiple clear Lexan panels that are
connected to four Delrin machined rails. The rails rest on the satellites launch rails and are the
sole point of contact to the satellite. Figure 58 shows this enclosure installed on the fully
83
assembled satellite, along with additional protective panels on the solar array wings – note that
the satellite is still compatible with the assembly jig stand whilst in the enclosure. Because the
solar array wings are externally protruding, their protective panels will remain installed up until
the satellite is mounted on the launch vehicle, even while being loaded into the deployment
system. They are outfitted with easy-to-remove thumbscrews and handles to facilitate the
removal at that stage. Meanwhile, the rest of the enclosure will be removed prior to integration
with the XPOD-Duo deployment system, and will then be shipped to the launch site, with
additional protective panels around the XPOD-Duo.
Figure 58: NORSAT-1 protective enclosure design
7.3
The large solar array wings on NORSAT-1 are a unique addition to the microsatellite. While the
rest of the structure is obtained from local machine shops, made in-house, or sourced from
commercial off-the-shelf (COTS) distributors, the solar array wings are manufactured by a third
party company who specializes in honeycomb composite materials for the space industry.
Because of the highly specialized field, relatively new technology, and amounts of qualification
testing required, these panels become an important economical consideration for the project. It
was decided that while two full sets of structure was procured as a “flight” and “spare”, only one
set of solar array wings would be procured, purely for economic reasons. Still needed, however,
was a set of geometrically similar wings to simulate these large solar arrays during various stages
of testing, in order to avoid damage to the expensive composite ones. The main design
requirements for these panels include:
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1) The Mock-Up Wings shall be similar to the intended flight wings in geometry.
2) The Mock-Up Wings shall be electrically similar to the intended flight wings.
While not fully weight representative, the mock-up wings serve as acceptable placeholders on
the satellite to achieve the correct geometry for various testing such as structural fit checks,
electromagnetic compatibility, uplink/downlink antenna, GPS antenna, VHF antenna, and
deployment testing. These mock-up wings are comprised of two thin aluminum sheets cut to the
correct geometry, and then sandwiched over spacers using epoxy adhesive to achieve the correct
thickness - the open edges are covered with Kapton tape to keep the panel easy to clean. They
are shown below in Figure 59.
Figure 59: NORSAT-1 mock-up wings (left), mock-up wings fitted on structure (right)
7.4
In order to characterize the performance of the communications subsystem design coupled with
the antenna placement on the satellite, and overall geometry for NORSAT-1, various antenna
pattern testing must be done in an anechoic chamber. To facilitate this testing, MGSE is needed
to help support the satellite in various orientations to point the antennas under test. Because this
GSE will be involved with the test itself and could potentially affect the results, some specific
requirements are derived for their design:
1) The Radio Frequency (RF) testing GSE shall be made of materials with a relative
permittivity (dielectric constant) that is as close to 1 (vacuum) as possible in order to
minimize its influence on test results
85
2) The RF testing GSE should be lightweight and easy to transport to off-site testing
facilities
3) The RF testing GSE design should leverage previous designs wherever possible to
minimize time, cost, and effort spent
4) The RF testing GSE shall provide four stable orientations at 45° increments around the
Y-axis – capable of being placed on a rotating platform without obstruction. This
provides all the necessary testing orientations for the antennas
The RF testing GSE will be used to perform testing on three different spacecraft antennas: The
S-Band Uplink/Downlink antennas, the GPS antenna, and the VHF antennas. Testing of these
three different antenna types will be performed in two separate off-site facilities – the test setup
and methods are further detailed in Section 8.5. In both facilities, a platform capable of rotating
360° is provided, on which the NORSAT-1 RF testing GSE is mounted. The GSE is composed
of a wooden baseplate that attaches to the rotating platform, on which is then stacked a large
polystyrene (EPS) block of foam to achieve a height similar to the source antenna in the facility.
Additional foam blocks are then stacked on top to achieve the appropriate orientations of the
satellite in 45° increments as seen in Figure 60. Each foam block is aligned and held in place by
using several wooden dowel pins, and the satellite rests directly on these foam blocks – wiring
for the antennas is routed through a hole at the center of the large foam block down to the
rotating platform. Note that the satellite is not fully integrated for these tests (only the outer
protruding elements are needed: the metallic structure, the LP Cassettes, and the antennas), and
is therefore relatively straightforward to re-orient to the required positions with the reduced
mass.
A large amount of this RF testing GSE was leveraged from previous SFL antenna pattern tests.
In fact, only the top foam blocks as seen in Figure 60 for the 45° and 90° orientations needed to
be slightly modified by the author to accommodate NORSAT-1. An additional foam block was
used under the 90° foam block in order to achieve the necessary clearance for the Langmuir
Probes. Note that the 45° GSE foam block doubles as the -45° orientation by simply rotating
the block under the satellite 180°.
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Figure 60: RF testing GSE blocks
Further detail and results of the specific RF testing that was done for NORSAT-1 using this GSE
is detailed in Section 8.5.
7.5
In the past, with the smaller sized GNB satellites (~7kg), the satellite could easily be picked up
via handles on one side of the spacecraft by a single person and manually loaded into its XPOD
deployment system with the aid of gravity and no additional GSE. NORSAT-1 is more than
double the mass and size of a GNB, so loading into the XPOD-Due in a similar way could not be
done in a reliable manor. Some design requirements for this deployment GSE are:
1) The XPOD loading GSE shall allow the spacecraft to be loaded into the XPOD and be
able to be removed once fully loaded in the XPOD
2) The XPOD loading GSE should double as the Deployment GSE since their purpose
entails fundamentally the same process
3) The Deployment GSE shall allow for horizontal deployment of the spacecraft from the
XPOD – Horizontal deployment minimizes risks associated with the large mass’ and
gravity
4) The Deployment GSE shall allow for full deployment of the VHF antennas
By adding a base plate and some retractable nylon ball transfers to the GSE support stand, the
satellite is able to slide freely on the support stand when needed. This provides a method to load
the satellite into the XPOD, and doubles as a method to perform deployment testing. This
deployment jig is designed to have all the necessary clearances to properly load into the XPOD-
Duo, and allow the XPOD door to fully open for a representative deployment. The height of the
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jig is sized to mate appropriately with the XPOD-Duo GSE, as shown in Figure 62, holding the
deployment system horizontally in order to perform this loading and deployment technique.
Commercial offset handles were also incorporated to easier handle the spacecraft and GSE
during this test. The deployment jig design is depicted below in Figure 61.
Figure 61: NORSAT-1 deployment jig design details
Figure 62: XPOD-Duo loading (XPOD-Duo GSE designed by Mike Ligori)
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8
All of the mechanically related integration and testing in which the author was involved for
NORSAT-1 is outlined in this section. Note that due to some schedule delays, not all of the
planned testing was complete at the time of writing this thesis.
8.1
Structural fit checks are a necessary step in the development process in order to avoid possible
surprises at a later stage in the spacecraft integration that could cause large delays. Upon
receiving all of the manufactured structural parts from the machine shop, each part is
individually inspected for any detrimental machining defects and errors. The parts are then fit
checked with their mating components and payloads in an effort to reveal any further defects in
the parts or design. The third step is to install all of the threaded inserts into the parts, in order to
allow a full spacecraft structural fit check to be performed, including all GSE parts (shown in
Figure 63). A path forward to resolve any defects found is determined at each step, and this
entire process is performed twice, for both the flight and spare set of structures. A list of the main
machining defects found during these fit checks is outlined below in Table 6, along with the path
that was taken for its rectification.
Table 6: List of main manufacturing defects found through inspection and fit checks
Defect Affected Rectification
+Y, -Y, -Z panels outer
dimensions slightly too large Their fit between the trays In-house filing
+Z tray had rounds machined on
the X panel mounting bosses Flush placement of the X panels In-house filing
Missing or wrong size holes in
+Z and –Z tray Mounting of various equipment
Returned to manufacturer for
modifications
Reaction wheel bracket machined
with some angled surfaces
Orthogonal placement of reaction
wheels
Returned to manufacturer for
rectification, use of copper shims
to adjust the height to account for
the missing material
Vent hole missing on risers, and
certain spots on the trays Sufficient venting of volumes In-house vent hole drilled
+X, -X, +Z, and solar cell
coupons were slightly warped
Installment of panels and
Laydown of solar cells on
coupon
Deemed acceptable and left as is
Protective panel holes all
misaligned
Installation of the protective
enclosure
Holes were enlarged in-house for
extra clearance
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Once the structural fit checks have been completed, and all errors rectified, the metallic parts are
sent for Iridite chemical conversion coating, in order to provide a layer of protection over the
bare aluminum surface.
Figure 63: NORSAT-1 structural fit checks, CLARA (engineering model) installed (left),
GSE enclosure installed (right)
8.2
SFL technicians build the flight and spare wiring harnesses for NORSAT-1. Once built, they are
each fit checked into the satellite structure to confirm the routing assumptions and lengths. In
many cases, the added length margin to the manufacturing drawings caused some of the
harnesses to be slightly longer than necessary, however this is not generally a problem – the
extra length can be accounted for by routing the wires in an “S” form between tie down points,
consuming much of the extra length. After the spare harnesses were built and integrated to the
Dirty-Sat (discussed in Section 8.3 Dirty-Sat Integration), certain lengths were updated for the
second flight build.
The larger harnesses, the Main and Payload ones, were slightly more complicated to fit check.
Due to the vast number of interconnecting connectors, and no easy way to sequence the order
that each connector be built in the harness, the resulting harnesses are seemingly tangled, causing
them to not fit in the satellite verbatim. During this fit check stage, the author proceeded to group
together common routed wires, while fit checking it into the structure. In some cases, this
involved removing crimped wires from certain simple connectors and reinserting them once
untangled in order to have them route as intended. Once untangled, and fully routed in the
structure, the harness is neatly bundled together in order to prevent it from intertwining, and to
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keep its form to facilitate its integration to the satellite during assembly. Figure 64 below shows
a before and after of this process for the Main wiring harness and Figure 65 shows the fit check
process for the Payload harness, showing the 3D model of the harness for comparison.
Figure 64: Freshly built (left) and untangled (right) flight Main wiring harness
Figure 65: Payload wiring harness, 3D model (left), fit check in structure (middle), and
untangled flight harness (right)
8.3
In order to mitigate much of the risk associated with a fairly new satellite design, a Dirty-Sat is
created as a fully functioning platform for testing. The Dirty-Sat is a close-to-fully representative
model of the flight spacecraft, using all flight representative (spare) equipment, including the
spare structure; however, it is assembled and integrated outside of a clean room (in a typical lab
environment) hence the name. This is done to make the integration and testing process simpler
and quicker to troubleshoot than performing everything in a clean room environment.
Structurally, the Dirty-Sat integration allows for a representative test and run-through of the
assembly procedure, allowing it to be further refined, and it also serves as a test-fit for the wiring
harness routing and design.
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Dirty-Sat integration for NORSAT-1 began at SFL in February 2015, was completed on March
12th
2015 and was followed by system level spacecraft functional testing. The integration was
performed semi-clean, in that parts were cleaned prior to integration, and gloves were used to
handle all equipment – this was done in order to reduce the risk of contaminating the spare parts.
Spare equipment was not however needed for all of the equipment, and alternatives were used
for the Dirty-Sat. Spare solar cells were not procured, and were simply not included in the Dirty-
Sat – instead, external power supplies are used to properly simulate them. The GSE Mock-up
Wings were used in the place of the solar array wings, and only one spare reaction wheel was
procured, therefore two flight reaction wheels were used in the Dirty-Sat. Some pictures of the
Dirty-Sat integration are shown in Figure 66 below. An engineering model (EM) of the battery
pack was also assembled for this integration.
Figure 66: NORSAT-1 Dirty-Sat integration
The Dirty-Sat integration and testing was able to reveal several insights to the satellite design
that would serve as refinements for the flight integration, these are briefly listed below:
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o The size of the Langmuir probe connectors on the electronics box were much larger than
anticipated, which affected the installation of the –Y panel. The payload provider has since
changed these connectors for flight.
o The data signals of the -Y sun sensor seemed to be “over-filtered” by the filtered MicroD
connector through the separation plate, causing the signals to be diminished or lost. These
data lines were since removed from the filtered MicroD connector, and alternatively
syphoned through the separation plate using individual, less intensive, feed through filters.
o The assembly procedure was refined, and in some cases modified. The integration of the
reaction wheels was accommodated at a very late stage in the integration in order to
minimize possible harm and contamination. This allowed for a much easier installation of the
wiring harness while connecting the trays together, providing extra room in the satellite to
route the wires without risk of damaging the wheels.
o Lessons were learned about the “Hand-Flex” coax cables that were used in the satellite. It
was found that they would get damaged quite easily when trying to form the path of the cable
once connected. Efforts were made to form the cable as best possible prior to installation to
avoid this. This is particularly of concern for the shorter cables (under 7 inches) that require
precise routing.
o The EMI gaskets used in various places in the satellite were found to fall out of their gasket
groove quite easily during installation. The gaskets were glued in place with very small
amounts of Room-Temperature Vulcanization (RTV) at their tips to avoid the gaskets falling
out during integration.
o The final location of all tie-down points for the wiring harness was confirmed. Many wire
lengths that were either slightly too short or too long were updated for improved routing in
the flight harness.
8.4
A large milestone for the NORSAT-1 project was electromagnetic compatibility (EMC) testing
of the Dirty-Sat. This testing began on March 16th
2015, was completed on March 24th
2015 and
was performed inside SFL’s in-house anechoic chamber (Figure 67). The purpose of this test is
to ensure that there is no significant interference that affects the performance of any of the on-
board electronics, and is particularly of importance due to the sensitivity of the AIS payload to
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very low frequency emissions, levels at which much of the spacecraft avionics is known to
produce noise. The effective signal strengths are measured at the receiver inputs of interest as
each of the spacecraft avionics are activated, in order to assess any produced interference. A
large external battery pack was used in order to fully simulate the power generation of the absent
solar cells; this ensures that the full functionality of the power system is active during testing.
Figure 67: NORSAT-1 EMC testing in SFL’s anechoic chamber
Overall, it was found that the NORSAT-1 platform provided a very “quiet” Radio Frequency
(RF) environment in the payload bay, exceeding the demanding requirement set forth by the AIS
receiver (#10 in Table 2). This successful result mitigated much of the risk of the novel design of
the platform, and confirmed the performance of the EMI reduction techniques that were
implemented in the structural design (Section 3.2.3).
8.5
In order to validate the spacecraft’s communication subsystem design and simulations, various
antenna pattern tests are done using spare structure of the satellite at various RF testing facilities.
8.5.1
S-Band and GPS antenna pattern testing was performed between June 17th
and June 19th
2015 at
the University of Toronto St. George Campus. The purpose of this testing is to characterize the
antenna performance, coupled with their mounting location on the full geometry of the spacecraft
bus. The tests involve measuring parameters of the spacecraft transmitted and received signals
94
with respect to a stationary source antenna inside an anechoic chamber. In order to achieve a
final spherical distribution of results around the satellite, measurements are taken while
performing 360° rotations of the spacecraft on an RF positioner platform, at four different tilt
angles of the spacecraft; details of these test orientations were discussed in Section 7.4. This
allows for the results to be plotted as a spherical distribution, and characterizes the
communication link of the satellite as a function of its orientation. A depiction of the test setup is
shown below in Figure 68. The results were post-processed by SFL’s Communications Engineer,
Clement Ma, and were all found to be acceptable.
Figure 68: Antenna pattern test setup
8.5.2
Because of the much lower frequency of operation, an alternate facility is needed to test the VHF
antennas. At the time of writing, the facility has not been finalized.
8.6
As per Requirement #14 in Table 2, all spacecraft deployable mechanisms must be tested in a 1g
environment. This requirement stems from a reliability standpoint – where if the deployment
mechanism is designed for a 1g environment, then it should certainly work in a 0g environment.
The main purpose of performing deployments tests of the satellite was to assess the performance
of the deployable VHF antennas. This became more necessary when the XPOD-Duo was slightly
modified to deploy horizontally instead of vertically. The areas of interest for this testing are to
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determine when the antennas start to deploy, the dynamics and path of their deployment, and
their behavior upon contact of a potential obstruction such as the launch vehicle.
The test setup for the deployment testing is shown in Figure 69. Note that the test setup has the
mounting legs of the XPOD facing upward, and will thus have the antennas deploying upward as
well. An engineering model (EM) version of the XPOD-Duo is used, along with the empty spare
spacecraft structure. The antenna hold down piece is attached to the XPOD pusher plate via
epoxy for these tests for convenience. The total mass of the spacecraft with the deployment GSE
for these tests is ~12kg. Large sheets of plastic are laid down as a track for the deployment in
order to reduce the friction as much as possible for a representative deployment.
Figure 69: Deployment testing test setup
8.6.1
Four deployments were performed in order to examine the behavior of the tape spring antennas
under various conditions. The tape springs are expected to stay stowed under the force of the
pusher plate upon deployment until the spring has reached its full length, however, a possible
scenario could exist whereby the antennas could slip out early due to the shock created when the
door opens, or due to the counteracting force of the tape springs themselves.
The first two tests are done in the configuration shown above in Figure 69, where the top of the
setup is all clear, allowing clear inspection of the antenna behavior from above. After these two
tests, it was confirmed that the antennas behave as expected, and do not move at all until the
spring has reached its full length. Figure 70 is a still shot during one of these deployments, at the
point just before the antennas start to deploy.
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Figure 70: Deployment test still shot, just before antennas deploy
For the subsequent two deployments tests, a large plate is placed atop the mounting feet of the
XPOD-Duo in order to simulate the mounting surface of the launch vehicle and examine the tape
spring behavior if contact is made. Due to lack of information of the exact mounting surface for
NORSAT-1 on its launch vehicle, the size of this “simulated launch deck” is based on an
expected mounting surface aboard a PSLV rocket for the NEMO-AM mission.
In the first of these two tests, the speed of the spacecraft deploying is fast enough that no contact
is made of the antennas on the plate, and the antennas are able to deploy unobstructed; a still shot
from this deployment is shown in Figure 71.
Figure 71: Deployment with mounting plate, no contact
For the second test, an extra 4kg is added to the spacecraft in order to slow down the
deployment, and ensure the antennas would contact the mounting plate; a still shot from this
deployment is shown in Figure 72 at the point of antenna contact. Both antennas sequentially hit
the mounting plate, causing them to bounce downward at first, and then continue to routinely
deploy.
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Figure 72: Deployment with plate, contact made
The speeds obtained during these deployment tests were ~1m/s, whereas in orbit, a much higher
speed of ~1.6m/s is expected. The lower speeds are caused by the various added frictional forces
in the test setup, largely due to gravity, and also due to the non-ideal condition of the EM XPOD-
Duo being used. As seen in the deployment shown in Figure 71, with a higher deployment speed,
the antennas have a better chance of deploying later, and avoiding potential obstructions.
However, in 0g, the deployment of the antennas themselves would also be much faster, because
they would not have to combat their own weight under gravity. Due to these numerous non-ideal
testing conditions, no clear conclusion can be made about the exact contact, if any, that would be
made by the antennas on the launch vehicle during deployment of NORSAT-1 in orbit; however,
this testing gives an idea of how the antennas could behave if contact is made, and can be used to
predict what kind of contact might be made once a clear mounting surface on the launch vehicle
is defined.
8.7
Due to the high sensitivity of the CLARA payload to particle contaminations, extra measures of
cleanliness are taken at all stages of the spacecraft integration. A bake-out of the flight structure,
prior to any avionics integration, was performed in July 2015 as an extra measure of cleanliness,
specifically for the CLARA payload. Due to having no electronic equipment in the flight
structure for this activity, a higher bake-out temperature can be achieved without risk of
damaging electronics, and a better result can be achieved. Included in the bake-out are all of the
flight structural components, mounting hardware, gasket material, and wire tie mounts. A bake-
out involves using heat in a vacuum environment in order to release particle contaminations from
something, as a method of cleaning, and forces materials to outgas at an accelerated rate. A “cold
98
wall” system is used; where-by a plate in a thermal vacuum chamber is kept extremely cold, and
allows a place for the removed particle contaminates to settle.
The bake-out was performed in the in-house large thermal vacuum chamber at the Space Flight
Laboratory, which is able to closely simulate space environments of a vacuum, and extremely
hot/cold temperatures. Seven temperature sensors are placed in various locations on the structure
in order to monitor and achieve a uniform temperature of interest. For this activity, a temperature
of 80°C was chosen based on the limits of the various included materials, and this was
maintained for approximately 38 hours. Figure 73 below shows the setup of this activity in the
large thermal vacuum chamber. Infrared lamps are directed towards the satellite in all directions
in order to obtain the high temperatures, and the satellite is suspended in the center via four steel
cables. The power of each lamp is controlled individually, allowing for some tuning to be done
while achieving a stable average temperature at each of the temperature sensors.
Figure 73: NORSAT-1 flight structure bake-out setup
Similar bake-outs will be performed at the component level of some of the other satellite
components, including the wiring harnesses, the solar panel wings, and the payloads. Finally, a
system level bake-out will also be performed on the integrated flight spacecraft with all avionics
prior to the CLARA payload integration.
99
8.8
Flight integration is currently on going for NORSAT-1 at SFL in the class 10,000 clean room. At
the time of writing, only the AIS receiver payload has been delivered and has been integrated
with the on-going assembly. Once the remaining flight payloads are received, they will be
immediately integrated in the flight satellite to complete the flight assembly in September 2015
(expected). Other outstanding items to be integrated include the power system, the spacecraft
panels and all solar arrays. Figure 74 below shows the current integration progress.
Figure 74: NORSAT-1 flight integration progress
A full assembly procedure for the battery pack has also been created by the author [21], and was
built by an SFL technician and engineer for flight. The fully integrated flight battery pack for
NORSAT-1 is shown below in Figure 75.
Figure 75: NORSAT-1 flight battery pack
100
9
NORSAT-1 represents Norway’s first scientific satellite, being a microsatellite carrying three
distinct and separately developed payloads. The main objectives of the payloads on-board are to
measure Total Solar Irradiance (TSI) levels from the sun, investigate space plasma
characteristics, and provide maritime ship tracking information through the Automated
Identification System (AIS). Much of the mechanical aspects of design, analysis, and testing for
the NORSAT-1 satellite bus has been completed and presented in this thesis. Subject to funding,
interest exists in developing subsequent multi-payload satellites using the NORSAT-1 design as
a platform; the development of NORSAT-2 has already begun, and is able to conserve
significant time, effort, and cost in the early stages by leveraging the presented design.
The structural design of NORSAT-1 has been developed from an early conceptual stage to a
detailed design currently under assembly and flight preparation. Verifications of all of the design
related requirements, including the driving requirements listed in Table 2, have been shown
through inspection or analysis, and in some cases, will be verified through testing in the near
future. The designs of two honeycomb composite solar array panels and two deployable whip
antennas mounted on the satellite have also been detailed to their current states. Various testing
and integration activities, including their test-specific mechanical ground support equipment
designs, are lastly presented with any significant results. Moving forward, several mechanical
aspects of development still exist for the NORSAT-1 satellite prior to launch. Completion of the
flight assembly, system level thermal vacuum testing, and system level vibration testing
represent some of these predominant tasks.
In thoroughly documenting the design and specific design decisions with their reasoning in this
thesis, the author has contributed a valuable reference for future microsatellite mechanical
designs. Efforts were made to design a very capable and modular satellite bus, for direct use on
future missions having similar requirements.
101
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[4] A. Fehlmann , "Metrology of Solar Irradiance," Zurich , 2011.
[5] J. D. Haigh, " Climate - Climate variability and the influence of the Sun," SCIENCE, vol.
294, pp. 2109-2111, 2001.
[6] P. Brekke and M. Osmundsen, "NORSAT-1: Total Solar Irradiance, Space Weather and
Ship Detection," in SOURCE Science Conference, Cocoa Beach, 2014.
[7] A. B. Donald A. Gurnett, Introduction to Plasma Physics: With Space and Laboratory
Applications, Cambridge University Press, 2005.
[8] A. Beattie, NORSAT-1 System Requirements and Verification Matrix v1.0, Toronto, ON:
SFL Internal Document # SFL-NS1-SYS-R001, 2014.
[9] S. Kanji, NORSAT-1 Structural Requirements and Verification Matrix, Toronto: SFL
Internal Document # SFL-NS1-STR-R001., 2013.
[10] S. Kanji and M. Chaumont, NORSAT-1 Assembly Procedure - Issue 0.4, SFL Internal
Document # SFL-NS1-MEC-G001, 2015.
[11] V. Tarantini, NORSAT-1 Thermal Design v2.0, Toronto: SFL Internal Document # SFL-
NS1-THM-D001, 2015.
[12] D. Diaconu, Mechanical Aspects of Design, Analysis and Testing of the Nanosatellite for
Earth Monitoring and Observation – Aerosol Monitor (NEMO-AM), Graduate Department
of Aerospace Science and Engineering University of Toronto, 2014.
102
[13] Hexcel Composites, HexWeb Honeycomb Sandwich Design Technology, 2000.
[14] T. N. Bitzer, Honeycomb Technology - Materials, Design, Manufacturing, Applications and
Testing, 1 ed., Springer, 1997.
[15] E. Greene, Failure Modes, Eric Greene Associates, 2013.
[16] S. Heimbs and M. Pein, "Failure Behaviour of Honeycomb Sandwich Corner Joints and
Inserts," Elsevier - Composite Structures, December 2008.
[17] Shur-Lok Corporation, Design Manual - Fasteners for Sandwich Structure, 1996.
[18] ECSS‐E‐HB‐32‐22 Working Group, Space Engineering - Insert Design Handbook,
Noordwijk: ESA Requirements and Standards Division, 2011.
[19] Structures and Mechanisms Division - European Space Research and Technology Centre,
Insert Design Handbook, European Space Agency, 1987.
[20] G. Bianchi, G. S. Aglietti and G. Richardson, "Optimization of Bolted Joints Connecting
Honeycomb Panels," in 1st CEAS - 10th European Conference on Spacecraft Structures,
Materials and Mechanical Testing, Berlin, 2007.
[21] S. Kanji, NORSAT-1 Battery Pack Assembly Procedure, Toronto: SFL Internal Document #
SFL-NS1-MEC-G002, 2015.
103
Figure 76: Insert strength as a function of core height for a 11mm diameter insert in a 3/16-
5052-.001 honeycomb core with aluminum face sheets under tension [19]
104
Figure 77: Insert strength as a function of core height for a 11mm diameter insert in a 3/16-
5052-.001 honeycomb core with aluminum face sheets under compression [19]
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