Assembly, Integration and Thermal Testing of the

Generic Nanosatellite Bus

by

Guy de Carufel

A thesis submitted in conformity with the requirements

for the degree of Master of Applied Science

Graduate Department of Aerospace Engineering

University of Toronto

© Copyright by Guy de Carufel 2009

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ABSTRACT

Title: Assembly, Integration and Thermal Testing of the Generic Nanosatellite Bus

Degree: Master of Applied Science

Conferred: 2009

Name: Guy de Carufel

Department: Aerospace Engineering

University: University of Toronto

This thesis describes the assembly and integration procedures, methods and strategies used for

the Generic Nanosatellite Bus (GNB) developed at the Space Flight Laboratory. The design of

the interconnection medium routing will be presented and aspects of thermal testing such as

thermal shock procedures and the satellite support structure design for the thermal vacuum

testing. The compliance of the assembly, integration and testing requirements is demonstrated

through validation and implementation. Step by step procedures are presented for GNB

assembly, solar cell bonding and thermal tape application. The evolution of the integration

design is described based on optimizing efforts and GNB design changes. Flexible circuits are

presented as an alternative to the conventional harness for future missions. Finally, general

assembly, integration and thermal testing recommendations are offered to add to the wealth of

knowledge acquired by SFL in the proper design of nanosatellites.

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ACKNOWLEDGEMENTS

My acknowledgments must start with Dr. Robert Zee. The amazing Canadian achievements

realized by your vision and creation of SFL should be recognized by all. Many thanks for

allowing me to be part of these historic accomplishments.

Special thanks to Cordell Grant for guiding me in my tasks and responsibilities. Your invaluable

experience and recommendations always put me back on the right track for reaching the next

step. The same must be said of Karan Sarda, with whom I had a great pleasure of working with.

Thank you Mihail Barbu for being understanding and accommodating for all the board layout

change requests I made. I would also like to thank Michael Greene for his great help with the

wiring harness construction and being a good friend and a source of personal motivation. I will

also thank Manuela Unterberger from the Austrian team who has not only given me invaluable

feedback on the documents sent but has also been a source of motivation in perfecting my work.

My experience at SFL has been not only valuable but pleasurable. I have my many new found

friends to thank for this great experience, and wish them all the best of luck in their endeavors.

My most sincere thanks go to my loving family who has always believed in me and supported all

my decisions, no mater how seemingly illogical they might have seemed.

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TABLE OF CONTENTS

ABSTRACT.................................................................................................................................. II

ACKNOWLEDGEMENTS .......................................................................................................III

TABLE OF CONTENTS ........................................................................................................... IV

LIST OF FIGURES ....................................................................................................................VI

LIST OF TABLES.................................................................................................................... VII

ACRONYMS AND ABBREVIATIONS................................................................................VIII

1.0 INTRODUCTION..................................................................................................... 1

1.1 The Space Flight Laboratory................................................................................... 1 1.2 Design Philosophy ..................................................................................................... 2 1.3 The CanX Program................................................................................................... 2 1.4 AISSat-1..................................................................................................................... 4

2.0 THE GENERIC NANOSATELLITE BUS ............................................................ 5

2.1 The GNB team........................................................................................................... 5 2.2 GNB Architecture ..................................................................................................... 6 2.3 Subsystems................................................................................................................. 7

2.3.1 Structure...................................................................................................................... 8 2.3.2 Thermal Control.......................................................................................................... 9 2.3.3 Power ........................................................................................................................ 10 2.3.4 Computer Hardware.................................................................................................. 11 2.3.5 Software and Communication protocols................................................................... 12 2.3.6 Attitude Determination and Control System............................................................. 13 2.3.7 Communication......................................................................................................... 14 2.3.8 Payload...................................................................................................................... 15

2.4 Deployment System................................................................................................. 19 2.5 Ground Support Equipment.................................................................................. 19

3.0 GNB ASSEMBLY................................................................................................... 20

3.1 Definitions................................................................................................................ 20 3.2 Assembly requirements .......................................................................................... 21 3.3 Assembly procedures and good practice............................................................... 22

3.3.1 GNB general assembly phases.................................................................................. 22 3.3.2 Flight ready cleanliness............................................................................................. 24 3.3.3 GNB Assembly procedures document...................................................................... 27 3.3.4 Assembly procedures revision history...................................................................... 32 3.3.5 Assembly process documentation............................................................................. 34 3.3.6 Assembly good practice............................................................................................ 35

3.4 Solar cell and S-Band patch antenna bonding procedures ................................. 37 3.4.1 Use of bonding compound........................................................................................ 37 3.4.2 Solar cell bonding procedures summary................................................................... 38 3.4.3 S-Band patch antenna bonding ................................................................................. 41

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3.4.4 Use of excessive RTV for AISSat-1 ......................................................................... 42 3.4.5 Solar cell inspection and functional testing .............................................................. 42

3.5 Thermal tape application ....................................................................................... 43 3.6 GNB ground support equipment........................................................................... 44

3.6.1 GSE requirements ..................................................................................................... 44 3.6.2 Early Design.............................................................................................................. 45 3.6.3 Final Design .............................................................................................................. 45 3.6.4 Materials used ........................................................................................................... 47 3.6.5 GSE Manufacturing .................................................................................................. 48

4.0 GNB INTEGRATION............................................................................................ 50

4.1 Integration Requirements ...................................................................................... 50 4.1.1 EMI reduction by twisted pairs................................................................................. 54

4.2 CanX-2 Integration Lessons................................................................................... 54 4.3 Integration Architecture ........................................................................................ 55 4.4 Integration design ................................................................................................... 56

4.4.1 Wiring harness components and materials ............................................................... 57 4.4.2 Harness routing and bridging strategy ...................................................................... 57 4.4.3 Mounting strategy ..................................................................................................... 59 4.4.4 Harness size verification ........................................................................................... 59 4.4.5 Coaxial cable routing ................................................................................................ 61 4.4.6 Computer board connector placement ...................................................................... 62 4.4.7 Integration procedures documentation...................................................................... 64

4.5 Wiring harness critical design changes................................................................. 66 4.5.1 Redundant use of OBC’s .......................................................................................... 66 4.5.2 Stage change of harness assembly ............................................................................ 67 4.5.3 Single to dual board design of CanX-4/-5 ................................................................ 68

4.6 Alternative Integration Concepts .......................................................................... 69 4.6.1 Flexible circuits......................................................................................................... 69

5.0 GNB THERMAL TESTING ................................................................................. 76

5.1 BRITE Solar cell Thermal Shock.......................................................................... 76 5.1.1 Solar cell T-Shock preparations................................................................................ 76 5.1.2 Solar cell T-Shock procedures .................................................................................. 77 5.1.3 Solar cell T-Shock results ......................................................................................... 78 5.1.4 Solar cell T-Shock test recommendation .................................................................. 79

5.2 Thermal-Vacuum Testing ...................................................................................... 79 5.2.1 CanX-2 jig structure modification ............................................................................ 80 5.2.2 BRITE and AISSat-1 TVAC structure ..................................................................... 81

6.0 AIT RECOMMENDATIONS AND CONCLUSION................. ......................... 86

6.1 General AIT Recommendations ............................................................................ 86 6.1.1 Assembly recommendations ..................................................................................... 86 6.1.2 Integration recommendations.................................................................................... 87 6.1.3 Thermal testing recommendations............................................................................ 88

6.2 Conclusions.............................................................................................................. 88 REFERENCES............................................................................................................................ 90

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LIST OF FIGURES

Figure 1: Exploded view of GNB architecture ............................................................................... 7 Figure 2: GNB Subsystems with associated components............................................................... 8 Figure 3: GNB Payload bay.......................................................................................................... 15 Figure 4: BRITE payload.............................................................................................................. 16 Figure 5: CanX-4/-5 payload and some of the optional components ........................................... 17 Figure 6: AISSat-1 Payload .......................................................................................................... 18 Figure 7: BRITE fit-check assembly with some corrected anomalies.......................................... 22 Figure 8: First and second test assemblies.................................................................................... 23 Figure 9: Cleaning steps for structures including hole and surface cleaning. .............................. 26 Figure 10: Clean structures of three complete sets for the BRITE satellite.................................. 26 Figure 11: Example of glue residue on wire................................................................................. 27 Figure 12: Procedures presentation system.................................................................................. 32 Figure 13: Example of commented changes in copy of revised document .................................. 34 Figure 14: Assembly tracking documentation example................................................................ 34 Figure 15: Mechanical hardware sorting and proper storage examples ...................................... 36 Figure 16: S-Band patch antenna bonding.................................................................................... 41 Figure 17: Solar cell short functional test ..................................................................................... 42 Figure 18: Thermal tape application techniques........................................................................... 43 Figure 19: Panel and magnetometer boom covered with gold thermal tape.................................43 Figure 20: Early designs of the GSE............................................................................................. 45 Figure 21: GSE Tray support rails and legs.................................................................................. 46 Figure 22: GSE panel support jigs ................................................................................................ 46 Figure 23: GSE Protective Enclosure ........................................................................................... 47 Figure 24: New GSE support legs variant .................................................................................... 47 Figure 25: GNB Integration Architecture ..................................................................................... 56 Figure 26: Photograph showing the bridging strategy used for the GNB wiring harness ........... 58 Figure 27: Maximum bundle height determination from CAD model ........................................59 Figure 28: Bundle size approximation representation.................................................................. 60 Figure 29: Harness inspection during panel assembly.................................................................. 61 Figure 30: Coaxial cable routing design for UHF and S-band (Left) and ISL (right) .................. 62 Figure 31: Board stack showing connector layouts for BRITE.................................................... 63 Figure 32: CanX-4/-5 payload harness over board stack with new dual board design................. 63 Figure 33: Tie-point representation system .................................................................................. 64 Figure 34: Routing table for all harnesses (for representation only – text not visible) ................ 65 Figure 35: Sun Sensor harness schematic..................................................................................... 66 Figure 36: Partial harness assembly to the +Z tray....................................................................... 67 Figure 37: Payload computers connector layouts (Left: BRITE, Center and Right: CanX-4/-5). 69 Figure 38: Modular (left) and daisy-chain (right) concepts for use with flexible circuits............ 73 Figure 39: Flexible circuits manufacturing capability examples.................................................. 74 Figure 40: Flexible circuit implementation strategy from first test version to flight version....... 75 Figure 41: T-Shock panel stack 1 (+X,-X,-Z) assembly............................................................... 77 Figure 42: Temperature profile of stack 1 during T-shock for 25 cycles ..................................... 78 Figure 43: Modified CanX-2 TVAC structure in YORK TVAC chamber .................................. 80 Figure 44: Suspension strategy for CanX-2 using cables and turnbuckles................................... 81

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Figure 45: Conceptual designs for the GNB TVAC support structure......................................... 83 Figure 46: Final design of the GNB TVAC support structure...................................................... 84 Figure 47: AISSat-1 TVAC structure shown in the DFL TV4 chamber ...................................... 85

LIST OF TABLES

Table 1: GNB design team.............................................................................................................. 5 Table 2: AIT Definitions............................................................................................................... 20 Table 3: Assembly Requirements ................................................................................................. 21 Table 4: GNB Assembly phases .................................................................................................. 24 Table 5: Cleaning procedures ....................................................................................................... 25 Table 5: Summary of GNB assembly procedures......................................................................... 28 Table 6: Guideline for when to take photographs during assembly ............................................. 31 Table 7: GNB Assembly Procedures document revision history ................................................. 33 Table 8: Good assembly practice.................................................................................................. 35 Table 10: GNB solar cell bonding procedures summary.............................................................. 38 Table 11: GSE Requirements........................................................................................................ 44 Table 12: GSE materials ............................................................................................................... 48 Table 13: Manufacturing recommendations ................................................................................ 49 Table 14: Integration requirements [27] ....................................................................................... 50 Table 15: AIT recommendations from CanX-2 development ...................................................... 54 Table 16: Benefits and limitations of flexible circuits for GNB integration ................................ 71 Table 17: BRITE solar cell T-shock procedures summary........................................................... 78 Table 18: GNB TVAC support structure requirements ................................................................ 82 Table 19: GNB TVAC support structure assembly order............................................................ 84

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ACRONYMS AND ABBREVIATIONS

ADC Analogue to Digital Converter ADCC Attitude Determination and Control Computer ADCS Attitude Determination and Control Subsystem AIS Automatic Identification System AIT Assembly Integration and Testing AWG American Wire Gage BRITE BRIght Target Explorer CAD Computer Aided Design CanX Canadian Advanced Nanospace eXperiment CNAPS Canadian Nanosatellite Advanced Propulsion System COTS Commercial Off-The-Shelf DET Direct Energy Transfer EMI Electro-Magnetic Interference GNB Generic Nanosatellite Bus GOC GNB Optical Camera GPS Global Positioning System GSE Ground Support Equipment HKC House Keeping Computer ISL Inter-Satellite L ink ISS Inter-Satellite Separation IR Infra Red LEO Low Earth Orbit LFFT Long Form Function Testing MOST M icrovariability and Oscillations of STars NASA National Aeronautics and Space Administration OBC On Board Computer PAYC Payload Computer PAY-POW Payload Power PCB Printed Computer Board PPT Peak Power Tracker RTV Room Temperature Vulcanizing RW Reaction Wheel SFL Space Flight Laboratory TVAC Thermal VACuum UHF Ultra High Frequency UTIAS University of Toronto Institute for Aerospace Studies VHF Very High Frequency WCC Worst Case Cold WCH Worst Case Hot

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1.0 INTRODUCTION

As an objective with this Master’s degree, the author wished to gain as much insight as possible into the

general design of a nanosatellite and the full life cycle of its development. His involvement with the

Assembly, Integration and Testing (AIT) of the Generic Nanosatellite Bus (GNB) was the perfect means

to achieve this objective. It required a general system level appreciation of the overall design of the

satellite as well as a thorough understanding of the subsystem integration aspects. The development of

the assembly and integration strategies and his involvement in different aspects of thermal testing gave

the author a direct experience in the different levels of the development cycle of a nanosatellite. The

author made a conscientious effort to present this thesis as a reference document for the AIT efforts of

future upcoming missions. The proper assembly, integration and testing of a satellite is of crucial

importance for the mission’s success. It is hoped that the author’s contributions have helped answer this

need for the BRITE, CanX-4/-5 and AISSat-1 missions.

This thesis will first introduce the Space Flight Laboratory (SFL) as well as the CanX program and the

AISSat-1 mission. It will then be followed by a description of the different subsystems of the GNB based

nanosatellites developed at SFL. Particular focus will be given to AIT aspects of each subsystem with

mention of the author’s contributions. The thesis will then be divided in three main Sections: assembly,

integration and thermal testing. The assembly Section will focus on the experiences gained through the

different assembly phases related to the BRITE nanosatellite mission. It makes use of this experience to

present a generic approach to the assembly process for the GNB as well as the design of its enabling

ground support equipment. The integration Section will begin with a treatment of the integration

requirements as well as lessons learned from past AIT efforts on the CanX-2 missions. The Section will

then cover general strategies for the design of the GNB wiring harness as well as significant GNB design

changes which had implications on integration. Potential future integration concepts are presented to be

considered for the next generation of the GNB architecture. Section 5 will cover work done related to the

thermal shock testing of the solar cells of the BRITE mission and the suspension frame design for

Thermal Vacuum testing of the AISSat-1 and BRITE missions. Finally, this thesis will present AIT

recommendations and conclusions.

1.1 The Space Flight Laboratory

The Space Flight Laboratory, or SFL for short, is led by Dr. Robert Zee, who founded SFL in 1998. SFL

offers an educational program at the University of Toronto Institute for Aerospace Studies (UTIAS) that

2

allows students to earn a masters degree in applied science in two years through hands on engineering

experience in the design and development of nanosatellites. Throughout the years, SFL has grown

considerably, and has expanded to include more full-time engineers that act as mentors for current

students. Currently the lab includes 13 students and 15 full-time engineers. The management

organization follows a project-department matrix system. This means that all students are assigned to

particular projects, or missions, and report to the project manager of that mission. In turn, each student is

assigned to a particular satellite function, which may span to include several projects. Each function is

lead by a full-time engineer who guides the students assigned to the function. SFL follows a particular

design philosophy which has resulted in a unique design approach, permitting this lab to distinguish itself

from other similar university programs involved in the development of nanosatellites and CubeSats.

1.2 Design Philosophy

At the Space Flight Laboratory, the design of satellites follows an approach based on the Microspace

Philosophy [1] in order to provide rapid access to space at a fraction of the cost of the traditional

approach. In general, this philosophy involves designing with reliability in mind, while building on

previous experience, avoiding unnecessary complexity and accepting a reasonable level of risk. Reducing

recurring costs by the use of generic components, as the use of the Generic Nanosatellite Bus (GNB), is

also practiced whenever possible. Also, making use of Commercial of the Shelf (COTS) components can

significantly reduce design cost and increase the reliability of the system. At SFL, space-grade

certification for purchased parts is not a requirement. Instead, the space worthiness of components is

validated through a set of qualification tests such as temperature, pressure and radiation testing in order to

simulate the space environment. Ultimately, this design philosophy enables other parties such as the

scientific community or commercial enterprises to have rapid access and cost effective means of

conducting meaningful science or testing innovative instruments in Low Earth Orbit (LEO) On-Board a

reliable satellite platform. This approach has been used for all missions at SFL, including the

Microvariability and Oscillations of Stars (MOST) mission, launched on June 30, 2003, for which SFL

was a key player.

1.3 The CanX Program

Following the experienced gained from the MOST mission; SFL commenced the Canadian Advanced

Nanospace eXperiment (CanX) in 2001. This program is a low-cost, rapid quick-turnaround program

involving Nanosatellites (satellites with a mass from 1 to 10kg) that conduct meaningful science and

3

space experiments. The first such mission was CanX-1 which was launched On-Board the same rocket as

the MOST satellite. This mission is considered a programmatic pathfinder, from which the CanX

program was initiated.

The second mission of the CanX program, CanX-2, was successfully launched from Sriharikota, India,

On-Board the PSLV-9 rocket in April 2008. The primary objective of this mission was to demonstrate

technologies to be flown on future missions. These include a cold gas propulsion system, the attitude

determination and control system, the communication system and the software operating system. This

mission also carried scientific payloads from universities across Canada including a GPS radio

occultation experiment, a spectrometry experiment and a material degradation experiment [2]. On-Board

the same rocket, the NTS (CanX-6) mission, involving a 20x20x20cm nanosatellite, was successfully

launched. This mission’s purpose is to demonstrate the ability to receive from space Automatic

Identifying System (AIS) transmissions from ships near the Canadian coasts, a mission for COM DEV

International Ltd.

Missions that are currently in development include CanX-3, and a dual satellite mission involving CanX-

4 and CanX-5. CanX-3, better known as the BRIght Target Explorer (BRITE) mission, is in the final

integration and testing phase in which the author is heavily involved. The BRITE mission is to study the

most luminous starts using differential photometry, which studies the relative intensity of different stars to

quantify stellar oscillations and better understand stellar life cycles [3]. BRITE will observe particular

star fields for periods from hours to months, which will require very precise and consistent pointing,

satisfied through the use of a star tracker. The BRITE mission is composed of at least two satellites,

called UniBRITE and BRITE-Austria. UniBRITE is under assembly at SFL, whereas BRITE-Austria

will be assembled in Austria to permit Austria to gain experience in the full assembly of a satellite, a first

for that country. Both BRITE satellites follow an identical design except for a difference in the optics

used in the On-Board telescope, such that each satellite is focused on a different light spectrum.

CanX-4/-5 is a dual identical nanosatellite formation flying demonstration mission which will make use

of cold gas technology developed in-house and flown on CanX-2, and demonstrate a formation flying

algorithm developed at UTIAS [4]. CanX-4/-5 will make use of advanced algorithms developed at

UTIAS to set new milestones in formation flying. Potential future uses for this developed technology

include remote sensing applications and on-orbit satellite servicing.

4

BRITE, CanX-4, and CanX-5 are all nanosatellites that follow the GNB architecture as a mean to reduce

recurring cost and accelerate the design process. The GNB design will be further explained in Section

2.0. The author’s work was mainly focused on the GNB Assembly, Integration and Testing (AIT) save

for the CanX-2 Thermal Vacuum (TVAC) support structure modification presented in Section 5.2.1,

which was also instrumental in the design of the GNB TVAC structure.

The CanX program has given the opportunity to over fifty graduate students to design, test, build and

operate nanosatellites; a very unique Canadian experience. This program has permitted SFL to gain

leading expertise in the field which accounts for the large amount of new missions at the laboratory.

1.4 AISSat-1

The AISSat-1 mission also makes use of the GNB system to support an Automatic Identifying System

(AIS) sensor developed in Norway. The objective of this mission is to investigate the performance of an

AIS sensor in low Earth orbit which would tracks maritime vessels in the Norwegian territorial waters.

As a secondary objective, the sensor would also triangulate the position of the transmitting vessels

through continued observations of their transmitted signal [5]. Minor adjustments to the standard GNB

platform will be required to accommodate the AIS sensor and its VHF antenna. The author’s

involvement in this mission includes its assembly and integration procedures development, the spacecraft

assembly, as well as its thermal vacuum testing.

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2.0 THE GENERIC NANOSATELLITE BUS

The Generic Nanosatellite Bus (GNB) is a spacecraft bus developed at SFL which was created to enable

the BRITE and CanX-4/-5 missions. In order to satisfy the power and size requirements of both missions,

a larger satellite bus compared to CanX-2 was required, which led to the final size of 20x20x20cm. The

cube form factor simplifies the thermal and power subsystems design. Since power, payload volume and

data processing requirements of both BRITE and CanX-4/-5 are very similar, it was logical to produce a

generic bus in order to reduce recurring costs. The idea behind the design of the GNB is to offer a

platform with standard components which could provide degree-level three-axis determination and

control, a communication link to ground, power distribution and switching, and data processing and

storing capabilities for a variety of payloads. Once the GNB is fully designed, quick access to space is

possible for any third party that wants to fly a payload that can be accommodated by the standard GNB

platform. Most components are optional and can be added if required for specific mission requirements.

The dual tray design of the GNB, which will be explained in Section 2.3.1, offers modular assembly, easy

access to different components and a mean for rapid assembly and disassembly of the spacecraft. From

the launch vehicle, it is deployed using the XPOD separation system which will be briefly explained in

Section 2.4.

2.1 The GNB team

Much effort has been invested in the development of the Generic Nanosatellite Bus since its conception.

Many student have been involved, some of which have continued as full-time engineers after the

completion of their master’s program. Table 1 lists all members who have been involved with the GNB

design, including the staff members, current masters students and a doctoral student who developed the

CanX-4/-5 formation flying algorithm.

Table 1: GNB design team

Name GNB Responsibilities Period Staff Members

Robert E. Zee SFL Program Manager, Design Review Se p 01 - Present Freddy Pranajaya Advanced Concept Manager, Nanosate llite Launch Service Sep 01 - Present

Cordell Grant CanX Project Manager, Structural Desi gn Sep 03 - Present Alex Beattie AISSat-1 Program Manager, Communicatio n Systems Sep 01 - Present Daniel Kekez Computer Hardware and Software Sep 03 - Present

Stephen Mauthe Propulsion Design and Ejection Syste m July 04 -Present Karan Sarda Structural Design and Thermal Control A ug 04 - Present

6

Mihail Barbu Computer Hardware Jul 07 - Present Tarun Tuli Computer Hardware Jan 05 - Present Manuchehr Ebrahemi Computer Hardware Oct 08 - Present

Nathan Orr Power Systems and Software Jul 05 - Pres ent Stuart Eagleson Attitude Control Aug 04 - Present Milko Dimitrov Communication Hardware Oct 08 - Pre sent Roman Ronge Software Development Oct 08 – Present

Masters Students Benoit Larouche Structural and Separation System Ma y 06 – Aug 08

Jonathan Grzymisch ADCS Hardware Jun 06 – Jul 08

Chris Short Orbital Simulations Sep 06 – Aug 08 Maria Short Ground Station Software Sep 06 – Aug 08 Adam Phillip ADCS Software Jun 06 – Apr 08

Guy de Carufel GNB AIT Jul 07 – Present Mark Dwyer Embedded Software Aug 07 – Present

Michael Greene ADCS Hardware Sep 07 – Present Amee Shaw Communications Hardware Sep 07 – Present Grant Bonin Power Hardware Sep 07 – Present Mohamed Ali Mechanical Testing Sep 07 – Present

Scott Armitage CanX-4/-5 AIT Sep 08 – Present Jenifer Elliott CanX-4/-5 Thermal Control Jul 08 – Present Jacob Lifshits GPS and Spacecraft Test Software Sep 08 – Present

Miru Choi Ground Station Software Sept 08 – Present Marc Fournier ACS Hardware and Software Sept 08 – P resent

Patrick Gavigan Computer Systems Jan 09 – Present Doctoral Student

Jesse Eyer CanX-4/-5 Formation Flying Algorithm Sep 05 – Present

2.2 GNB Architecture

The GNB architecture follows a 20x20x20 dual tray design, with a central payload bay between trays.

Most components are mounted directly to either tray, with the exception of the magnetometer boom, the

various antennas, most sun sensors, the magnetorquers and solar cells. The GNB includes a bank of

generic components which may or may not be present on each mission, based on mission specific

requirement. Figure 1 shows the layout of all components in the GNB, with mention of optional

components and their associated missions.

The position of each component both internally and externally, was dictated by several constraints which

led to the current configurations. As an example, the S-Band patches must be on opposite panels and

different from the inter-satellite patches for CanX-4/-5, which lead to their final placement to the Z

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panels. For efficient use of volume and simplicity in assembly and harness integration the computer

boards were grouped and stacked in the +Z tray. The location of the dedicated volume for the payload

was centralized within the satellite in order to reduce the distance between the center of mass with the

center of volume. Also for many applications such as the propulsion system in CanX-4/-5, it is desirable

to have the payload aligned with the mass center to avoid undesired moments from thrusting or

environmental forces. The architecture of the GNB was designed by Cordell Grant and Benoit Larouche.

Full justification to the GNB layout can be found in [6].

Figure 1: Exploded view of GNB architecture

2.3 Subsystems

The GNB architecture is divided into eight subsystems: structure, thermal control, power, computer

hardware, software, communication, Attitude Determination and Control Systems (ADCS) and payload.

Figure 2 shows the tree of subsystems with their associated components. Note that some components

such as the GPS receiver, Inter-Satellite Link Radio, Rate sensors, VHF beacon and imager are optional

and mission dependent. Also, the amount of solar cells used depends on surface area available on panels

8

and the mission specific power requirements. The following subsections will describe in greater detail

each subsystem and its associated integration considerations.

Figure 2: GNB Subsystems with associated components

2.3.1 Structure

The structural design of the GNB has evolved into the current dual tray design, with a standard volume

payload bay between both trays. The structural strength of the satellite responsible for supporting the

vibration loads experienced during launch is mainly ensured by the trays and the payload support

structure. Depending on the mission, the trays are made of either aluminum or magnesium and are nickel

plated for galvanic corrosion resistance and improved conductivity for grounding purposes. Openings in

the sides of the tray allow routing access for the wiring harness and permits access to the different

connectors on the various boards. Launch rails directly incorporated into the tray structures are used to

guide the satellite out of the XPOD deployment system during ejection. The 2mm thick panels with cross

braces offer additional rigidity to the structure and provides a mounting surface for the solar cells, most

sun sensors, the magnetometer boom, the magnetorquers and the various antennas. All other components,

such as the reaction wheels, various computer boards, battery assemblies, radios and some of the sun

sensors are mounted directly to either tray by the use of mounting bosses. The BRITE and CanX-4/-5

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structure was designed by Cordell Grant [6], while the stress and vibration analysis was conducted by

Benoit Larouche [7]. The AISSat-1 structure was developed by Stephen Mauthe.

2.3.1.1 Structural AIT aspects

For proper wiring harness integration, the tie-down points on the structure had to be defined earlier on

before its manufacturing. The author was responsible for revising the structural design such that it could

accommodate the wiring harness and allow proper assembly of its different components. Tie-down points

for the harness are secured using tie-wraps which make use of dedicated holes in the structure. The tie-

down point strategy and integration approach for BRITE, CanX-4/-5 and AISSat-1 developed by the

author is explained in greater detailed in Section 4.4.3.

2.3.2 Thermal Control

The thermal control of the spacecraft is mostly accomplished using passive methods. This includes

thermal tapes on the panel surfaces to achieve desired thermal balance between heat emission and

absorption as well as regulating heat conduction within the satellite with the appropriate choice of

component mounting strategies. A battery heater is also included to keep the batteries sufficiently warm

so that they are not damaged during charging. Since the thermal design is highly dependent on the final

orbit of the satellite, the design approach taken is to design two separate thermal tape strategies for both

baseline sun-synchronous orbits; dawn-dusk and noon midnight. Each baseline orbit is analyzed using

conservation of energy principles under Worst Case Hot (WCH) and Worst Case Cold (WCC) conditions

so that component temperatures are within requirements for all situations. The thermal design for all

three GNB missions was developed by Karan Sarda. Detail on the thermal design and analysis can be

found in the internal document [6].

2.3.2.1 Thermal AIT aspects

The author’s involvement with the thermal subsystem included the thermal tape application process

during the BRITE satellite assembly as explained in Section 3.5. Also, Thermal testing will be conducted

on the fully integrated satellite to simulate the space environment. At the unit level, a thermal shock test

(T-shock) is done on all electronic components and solar cells. The author was responsible for the T-

shock test of the solar cells, in order to validate their workmanship and demonstrate their survivability

under the extreme space temperatures and high temperature change rates. On the systems level, the fully

integrated satellite undergoes testing in a Thermal Vacuum (TVAC) chamber where the satellite

functionality is tested in the expected space environment. The TVAC testing is conducted to validate the

10

expected thermal distributions for both WCH and WCC for both baseline orbits. The author’s

involvement with the TVAC testing included the design of the TVAC chamber satellite suspension frame

for both BRITE and AISSat-1 as well as modifications to the CanX-2 suspension frame. Also, the author

is expected to assist the TVAC testing for the AISSat-1 mission. Thermal testing will be covered in

Section 5.0.

2.3.3 Power

The Power subsystem is responsible for the power generation, storage, regulation, control and

distribution. GNB Generic components for the power subsystem include the solar cells, the batteries with

their associated BCDR, the power board and the wiring harness. The payload computers, which are

mission specific, are responsible for power regulation and control of mission specific payloads and

components. Also related to the power subsystem are the separation switches which are magnetic

switches that are closed when the satellite is integrated in its deployment system, the XPOD system.

After the satellite is ejected, the switches revert to their nominal open position, which turns the satellite on

by allowing unregulated power supply to the power board. More information on the deployment system

will be found in Section 2.4. The power subsystem has been developed mainly through the efforts of

Nathan Orr, Mihail Barbu and Grant Bonin.

Some of the design for the power subsystem is derived from the CanX-2 mission; however, key

differences are present. The main difference being that the CanX-2 satellite made use of Direct Energy

Transfer (DET), where the bus voltage is governed by the battery voltage, whereas the GNB has the

option to use a Peak Power Transfer (PPT) architecture. Although PPT is more complex, it is more

efficient in that it keeps the power generated by the solar cells at its peak by voltage regulation at the

batteries. The PPT system is made possible by the use of the Battery Charge Discharge Regulator

(BCDR), which keeps the bus voltage at the peak solar cell output power. The GNB architecture includes

two batteries, each with an associated BCDR. Nominally, only one Battery/BCDR is used, with the

second as a cold spare. This design follows a modular approach, were both BCDRs can perform the same

function independently. Included in the PPT system are blocking diodes accomplished by ideal diode

ICs, which are used to prevent batteries from discharging through the solar cells during eclipse.

The power provided to most GNB component is regulated and controlled On-Board the power board and

distributed through its own distinct connector. Some components, such as the reaction wheel perform

their own power regulation. Power regulation On-Board the power board is accomplished by switch-

11

mode voltage regulators which can step-up or step-down the bus voltage to the required voltage for each

component. The power board is also responsible for over-current protection, which is accomplished at

both the software and hardware level with current sensor ICs and a 12-bit ADC. The power controller,

also On-Board the power board, makes use of a low power FPGA to accomplish power switching. The

power board also includes a test port which includes a satellite jumper ON/OFF to mimic the separation

switch, allows charging of the batteries from an external source and allows monitoring of the battery

voltage. More detail on the power subsystem can be found in [8].

2.3.3.1 Power subsystem AIT aspects

The power distribution makes use of Hirose DF-11 connectors and Teflon insulated copper braided wires.

These components are the distribution medium for the integration of all components in the GNB and will

be discussed in greater detail in the Section 4.4.1. The author was also involved with the BRITE power-

sub system for the assembly of its components, including the bonding of the solar cells to the panels

described in Section 3.4. For more efficient integration, the power board connector layout was

completely revised by the author for better harness routing. The Power board layout evolution is

described in Section 4.4.4.

2.3.4 Computer Hardware

The computer hardware is composed of two On-Board Computers (OBC), called the House Keeping

Computer (HKC) and the Attitude Determination and Control Computer (ADCC), and possibly a third

OBC, the Payload Computer (PAYC). The HKC is responsible for the overall housekeeping of the

spacecraft such as telemetry collection and logging, communication relay between the satellite and the

radio system and issuing commands to the power board for power control. The ADCC is responsible for

the Attitude Determination and Control System (ADCS) on board the satellite. This includes processing

and executing the commands received related to ADCS components as well as performing automated

attitude control. Both OBC’s have the capability to interchange their roles such that their responsibilities

can be reestablished in the event of malfunctioning of either computer. This redundant architecture is

ensured through splicing in the harnesses or dedicated wires. The PAYC is responsible for any telemetry

or command processing related to the On-Board payloads. For the BRITE mission, this is accomplished

by a single board which controls both power and telemetry for the BRITE telescope. The BRITE Payload

Computer (PAYC) is also referred to as the Instrument Computer (IC). For CanX-4/-5, a dual board

system is used, composed of a PAYC which is in fact an OBC identical to the HKC and ADCC in

hardware. A second daughter board called the Payload Power (PAY-POW) board is responsible for the

12

power regulation for the payloads such as the CNAPS propulsion system and the Inter-satellite Separation

System (ISS), and the PAYC. For AISSat-1, a single PAYC as in CanX-4/-5 is used. The OBC was

developed by Tarun Tuli, while the BRITE PAYC was developed by Mihail Barbu. Further details of the

hardware designs can be found in [9].

2.3.4.1 Computer Hardware AIT aspects

The computer board layouts went through many different iterations, in which the author contributed in the

connector placements such that their layout be well located for an efficient harness routing from computer

to components. The computer board connector layout designs will be covered in Section 4.4.6. The

implementation of the redundant functionality of the OBC impacted the wiring harness design by adding

a significant number of wires to include the connection to the other OBC. This will be discussed in

Section 4.5.1. The payload computer for the CanX-4/-5 mission was initially a single board design called

the Formation Flying Computer (FFC). This was changed to a dual board design. The harness

implications from the change from a single board to dual board system will be covered in Section 4.5.3.

2.3.5 Software and Communication protocols

The Software subsystem of the GNB nanosatellites is responsible for all command processing and

execution for all the GNB components. The software architecture is separated into two levels; the

bootloader and the application codes. The programming instructions for the bootloader reside on the

External EPROM of the OBC’s, while the application code resides in the processor’s internal FLASH and

external SRAM. The bootloader is responsible for the initialization of the UHF receiver (RX) and S-

Band transmitter (TX), as well as certain hardware initializations, such as the communication drivers

necessary for inter-computer and inter-thread communications. A transmitter select (TX select) governs

which medium the bootloader should respond to, either the Test port or the S-Band TX. Every instruction

for the bootloader is pre-loaded before flight and is protected by an overwriting protection such that it

may not be modified during flight. The application code makes use of the CANOE Operating System

(OS) which was developed in house by Tarun Tulli [10] and Cecilia Mok [11]. CANOE makes use of a

file system to keep track of used areas of the memory and assign space for new files. It also keeps an

array of Process Control Blocks (PCB), each containing a programming thread, thread state, a specific

Process ID (PID) and pointers to memory buffers. Through the driver threads on CANOE, any thread can

send a message to any other thread through these message buffers thereby activating these new threads.

A task scheduler allows the appropriate cycling through active threads for rapid execution of the different

commands received from the ground station or other threads. All communications to the satellites is first

13

processed by the House Keeping Computer (HKC) and then routed to the appropriate destination by the

communication drivers. More details can be found in the software architecture internal document [12].

All communications On-Board the satellites follow the NanoSatellite Protocol (NSP), derived from the

Simple Serial Protocol (SSP) used for the MOST mission, and was developed by Cecilia Mok [11] and

Daniel Kekez at SFL. The NSP protocol offers half-duplex or full-duplex communication. A Cyclic

Redundancy Check (CRC) is used during transmissions in NSP in order to detect corrupted packets. NSP

is used by the on-board software for inter-processor and inter-thread communications. More information

on the NSP protocol may be found in [13].

2.3.5.1 Software AIT aspects

For proper performance of the On-Board software, all programmed threads must be verified and tested

when integrated to the different components. The validation of all software is first done on what is called

the FlatSat, which consists of all integrated components without the satellite structure. The final

validation of the software is performed during TVAC testing, where the fully integrated satellite is

subjected to the different expected operating conditions in a space like environment. The satellite

undergoes Long Form Functional Testing (LFFT) under both a Cold Start and Hot Start condition. The

author will act as a TVAC operator to conduct the LFFT for the AISSat-1 satellite, which is being

developed by Jacob Lifshits.

2.3.6 Attitude Determination and Control System

The Attitude Determination and Control System (ADCS) is responsible for the pointing of the satellite to

satisfy payload specific requirements. The attitude determination of the satellite is ensured by various

sensors including six coarse and fine sun sensors, a 3-axis magnetometer and a 3-axis rate sensor in the

case of AISSAT-1 and CanX-4/-5 for body rate determination in the event of an eclipse. For BRITE, a

star tracker is used to provide very fine pointing determination in order for the telescope to always point

at the same star for prolonged periods of time. The sensor measurements are used by the On-orbit

Attitude System Software (OASYS), developed by Stuart Eagleson [14], in order to determine the

satellite attitude. OASYS then compares the actual satellite attitude with the desired attitude in order to

determine the required correction torques. These torques are produced by the attitude control system. A

set of three orthogonal reaction wheels residing in the –Z tray are used for fine attitude corrections. For

coarse corrections and momentum dumping to prevent the wheels from reaching saturation, a set of coils,

14

called magnetorquers, are used. The magnetorquers are mounted on three orthogonal panels, specifically

the +X, +Y and –Z panels. More details on the GNB ADCS design can be found in [15].

2.3.6.1 ADCS AIT aspects

As for all components on the GNB, the ADCS component harness routing design was performed by the

author. Particularly challenging was the routing of the sun sensors as they are widely distributed around

the satellite. The sun sensor harness routing documentation system will be described in Section 4.4.7. An

assembly fit check test performed by the author, as will be described in Section 3.3.1, led to corrections

being made for the magnetorquer by Michael Greene, so that the panels could be properly assembled to

the rest of the BRITE satellite.

2.3.7 Communication

This subsystem is responsible for telemetry transfer between the ground and the satellites and also

between satellites in the case of CanX-4/-5. This is a full-duplex radio communication system meaning

that the satellite and ground station can both upload and download data simultaneously. This is achieved

by UHF uplink (from ground to satellite) and S-Band downlink (from the satellite to ground). The UHF

receiver (∼400MHz) On-Board the GNB satellites can receive data at a rate of 4Kbps. This is achieved

by a quad-canted monopole antenna configuration which achieves a near omni-directional gain pattern.

The downlink is achieved by an S-Band transmitter (∼2200MHz) with a variable data rate (8-256Kbps)

and a variable modulation strategy (Binary or Quad Phase Shift Keying, BPSK or QPSK). These features

are used to increase data rates if the link is sufficiently strong. The S-band transmitter makes use of two

S-Band patch antennas, one on each Z panel. Since both uplink and downlink can function in any

attitude, no ADCS constraints are imposed by the communication system. For CanX-4/-5 a VHF beacon

and an Inter Satellite Link (ISL) radio is included. The beacon is used as a simple means to acquire

spacecraft health and tracking, thereby complementing the primary link system. The S-Band ISL radio

will by used by both CanX-4 and 5 for inter satellite navigation data exchange, which is a necessary

component for successful formation flying. Two patch antennas, similar to the downlink S-band patch

antennas, are bonded to the X panels. This system is also omni-directional. Details on the ISL can be

found in [16] and details on the UHF, S-Band and VHF radios can be found in [17]. The communication

systems for the GNB missions have been developed by Alex Beattie and Daniel Kekez.

15

2.3.7.1 Communication AIT aspects

For proper circular polarization of the communication signals in the case of the UHF, it is required that all

four antennas have a matching signal delay with 90° phase increments in a circular order. This requires

that all four coaxial cables running from the antenna connector to the UHF radio receiver be of equal

lengths. The same requirement is present for the S-Band and the ISL radio, such that the intersecting

region of radio coverage by both patch antennas is synchronized. The routing strategy for these coaxial

cables developed by the author will be explained in Section 4.4.5. The author has also developed the S-

Band patch antenna bonding procedures, as discussed in Section 3.4.3.

2.3.8 Payload

As shown in Figure 3, the payload bay occupies a central volume of 13x17x8cm between both trays. Any

external appendages such as the antenna for AISSat-1 are not included in this allowed volume. The

payload is always supported by the trays rather than the panels. It is possible that the payload includes a

structural frame which provides the required structural strength between both trays. If not, additional

structural braces are added between the trays, as is the case for AISSat-1. The payload may require

specific keep out zones where objects, such as the wiring harness, must not be present. The payload

communicates with the satellite bus through a dedicated payload computer situated at the bottom of the

OBC stack in the plus Z tray. Thermal control of the payload is usually achieved by passive methods,

such as a proper mounting strategy for heat conduction to the trays.

Figure 3: GNB Payload bay

16

2.3.8.1 BRITE Payload

BRITE makes use of a telescope developed in-house in order to perform its scientific mission of

differential photometry. This telescope is designed as a photometer, with a wide field of view (about 25

degrees) and a short focal length. The telescope contains no moving parts since its optics is designed for

a particular light spectrum. The mechanical design of the telescope has been developed by Cordell Grant

[18]. For its detector, much in house testing by Normand Deschamps, and subsequent validation by Mark

Dwyer, has led to the selection of the Kodak-KAI-11002 CCD imager [19]. Due to the stringent pointing

requirements to maintain the payload pointed at the same star over prolonged durations, a star tracker is

also included in the payload bay. The telescope CCD is operated by the Payload Computer, whereas the

star tracker is controlled by the ADCC. Figure 4 shows the complete payload for BRITE, including both

the telescope and the star tracker which are supported by the payload support structure. The telescope has

a dedicated computer board called the Payload Computer (PAYC).

Figure 4: BRITE payload

Cleanliness is of particular importance to this mission as any impurities could obstruct its optical

instruments. Cleanliness procedures are covered in Section 3.3.2. The assembly order and wiring

strategy of the BRITE satellite was highly dependent on the mechanical design of the BRITE payload as

will be covered in both the GNB assembly and integration Sections. The PAYC connector location

justification will be covered in Section 4.4.4. The harness between the telescope and the PAYC makes

use of a micro-D connector at the CCD end, and a DF-11 connector at the PAYC end. A coaxial cable

with SMA connectors is also used. More on harness connectors and wire types will be explained in

Section 4.4.1.

17

2.3.8.2 CanX-4/-5 Payload

The main payload for the CanX-4/-5 mission is the Canadian Nanosatellite Advanced Propulsion System

(CNAPS). This propulsion system will allow both CanX-4 and CanX-5 to perform formation flying.

CNAPS will provide the required thrust to both transfer between various formations and allow orbit

keeping. CNAPS makes use of liquefied sulfur hexafluoride (SF6) as its propellant and will make use of

results and experiences gained on its precursor, the NANOPS system that was used on the CanX-2

mission. It makes use of four independently controlled thrusting valves in order to correct any mass

center misalignments through active control or on-orbit calibration. The CNAPS system was developed

by Stephen Mauthe [20]. During ejection and commissioning of the satellites, both CanX-4 and 5 must

remain attached so that they do not drift apart and require valuable fuel to close the distance. The

component that will keep both satellites attached during commissioning is called the Intersatellite

Separation System (ISS) which was developed by Cordell Grant and Benoit Larouche [21]. CNAPS, the

GPS receiver and the Intersatellite link radio will communicate with the Payload Computer (PAYC),

which is in fact identical in hardware to the OBCs. The formation flying algorithm used on the PAYC

was developed by Jesse Eyer [22]. This algorithm requires communication between the two satellites

which is accomplished by the InterSatellite Link (ISL) radio. As optional GNB components, CanX-4/-5

makes use of a 3-axis rate sensor, a GPS antenna and receiver for accurate position determination and the

GNB Optical Camera (GOC) imager which will take color photographs of the separation process. The

preliminary design of the GOC imager was developed by Norman Deschamps [19]. The power regulation

for CNAPS, the ISL, the GPS receiver, the ISS and the Payload Computer is done on the Payload Power

(PAY-POW) daughter board.

Figure 5: CanX-4/-5 payload and some of the optional components

18

In order to make the wiring harness generic, the reference mission used by the author was CanX-4/-5 as it

contains all possible GNB components. The BRITE’s harness is identical, save for wire subtractions for

components it does not have and the differences in the payload harness. During the design process, the

payload board changed from a single board to a dual board configuration, which affected the integration

strategy. This will be explained in Section 4.5.3.

2.3.8.3 AISSat-1 Payload

The AISSat-1 payload provided by Norway includes an Automatic Identification System (AIS) sensor

and a VHF antenna which will receive AIS signals from ships in Norwegian waters. Because of the extra

structural strength required to support the antenna, the tray design was slightly altered from the standard

GNB tray in order to support the antenna mounting bracket. Also, the payload box holding the AIS

sensor is supported to the trays by the use of support brackets. The dedicated computer board for the

payload is the Payload Computer (PAYC), which in hardware is identical to the other OBCs. The

payload is powered directly through the Power Board (power board), hence does not require a daughter

payload power board as is the case for CanX-4/-5. As optional components, AISSat-1 will include a GPS

receiver and antenna and a 3-axis rate sensor. More on the AISSat-1 payload can be found in [5].

Figure 6: AISSat-1 Payload

19

For integration, the harness for all generic components in this satellite is identical to the other two

missions. Only the payload specific harness and its routing strategy are peculiar to AISSat-1. The

Thermal Vacuum (TVAC) support structure made use of this mission as its reference mission since it

must accommodate the extra pre-deployed antenna boom, not present on the other missions. The TVAC

structure will be presented in Section 5.2.2.

2.4 Deployment System

The deployment system for the GNB based satellites is called the XPOD. It was developed by William

O’Brien, Daniel Kekez, Stephen Mauthe and Freddy Pranajaya. It follows a “Jack in the box” approach

where a single spring ejects the satellite once the door is opened by the use of a pusher plate. The door

makes use of a release mechanism that consists of a clamping mechanism which is preloaded to open and

kept closed by the use of a Vectran cord. When the ejection command is given, the cord is broken by the

use of voltage applied to a highly resistive wire, thereby burning the cord. Since the CanX-4/-5 satellites

will be joined when ejected, an XPOD Duo, which can deploy both satellites, will be used. More

information on the deployment system can be found in [23].

2.5 Ground Support Equipment

In order to protect the satellite during its storing and assembly and for protection of the various

components during its manipulation, the satellites make use of Ground Support Equipment (GSE) called

the protective enclosure. For the GNB protective enclosure, initial designs where proposed by Benoit

Larouche [7]. The author made further iterations to the design and manufactured in house the final

Ground Support Equipment to be used for the GNB missions. The complete details on the design

evolution and manufacturing are presented in Section 3.6.

20

3.0 GNB ASSEMBLY

This section will cover the assembly aspects under consideration by the author in regards to the GNB

based missions, with particular emphasis on the BRITE mission. After defining the different aspects of

AIT, this section will cover requirements related to assembly. The work presented here was strongly

influenced by the AIT approaches used for the CanX-2 mission. Assembly strategies verified through

actual satellite assembly will be presented, including good assembly practice, cleanliness measures, solar

cell bonding procedures, thermal tape application measures and a description of the GSE protective

enclosure for the GNB. The author’s responsibilities concerning assembly consisted of developing the

assembly procedures for the GNB missions and performing the physical assembly of the BRITE satellite.

3.1 Definitions

In an effort to describe the different aspects of the AIT of a satellite, the following definitions are

suggested in Table 2.

Table 2: AIT Definitions

Assembly The mechanical action by which different components are assembled to form larger

subsets which in turn are assembled to form the complete spacecraft.

Integration The process by which the spacecraft, composed of a collection of assembled

components, is made functional through the use of an interconnection medium.

Testing The process by which the confidence of the satellite to properly operate in orbit is

satisfied by environmental and functional validation.

Although integration can be somewhat synonymous to assembly, integration in this thesis describes the

act of turning the assembled satellite functional by interconnecting the different components and control

hardware through the use of the wiring harness. Hence, any harness related assembly activities will be

covered in the integration Section. In order to perform system level testing as opposed to unit level

testing, a fully integrated satellite is required. Integration will be covered in Section 4.0 and testing in

Section 5.0. The Ground Support Equipment (GSE), presented in Section 3.6, is considered a tool for

assembly. Assembly also involves the correct handling practice, cleanliness considerations and

inspection, all of which will be discussed in greater detail in Section 3.3.

21

3.2 Assembly requirements

In order to properly undertake the task of assembling the spacecraft, a complete understanding of the

requirements is necessary. Table 3 shows requirements originating from the GNB systems requirement

document [24], for which every aspect of the satellite design, including AIT, must comply. The table is

divided into general systems and AIT specific requirements.

Table 3: Assembly Requirements

General systems requirements [24]

Sys 1.8 The BRITE and CanX-4/-5 assembly procedures shall be documented such that a skilled

and experienced team could build an exact duplicate of any of these satellites.

Sys 2.3 The mass of each satellite using the GNB shall not exceed 7kg. Mass margins shall be no

less than 30% in the preliminary design phase, and no less than 20% in the detailed design

phase.

Assembly specific requirements [24]

Sys 9.1 The spacecraft design shall permit partial assembly and disassembly.

Sys 9.2 The spacecraft design shall allow complete assembly and/or disassembly with simple tools.

Simple tools include standard screwdrivers, tweezers, pliers, wrenches and cutters.

Sys 9.3 Ground Support equipment (GSE) shall be created to facilitate assembly, protect the

satellite at all stages of assembly, and protect the more delicate satellite components (solar

panels, sun sensors, etc.) when the satellite is disassembled.

Sys 9.4 If pre-deployed UHF or VHF antennas are used, they should be removable from the fully

assembled satellite without any additional disassembly of the satellite.

Sys 9.5 Electrical connectors, fasteners, and any other rigid connections shall be staked with RTV

(eg. CV 1142) to prevent loosening and possible disconnection during launch.

The general systems requirements listed are pertinent to AIT since proper documentation of assembly

processes used is necessary for repeatability. Also, documenting procedures ensures that no tasks are

omitted when followed. In this respect, the author produced a document on the complete GNB assembly

procedures [25] and on the solar cell bonding procedures [26]. Also underlined by these requirements is

the necessity to properly track changes in these documents and maintain good version control. To satisfy

requirement 2.3, the mass of every component to be assembled was weighted on a microbalance and

recorded, such that the actual mass of the satellite is confirmed as less than 7kg. The AIT specific

requirements mentioned entail an efficient and reliable assembly of the satellite, such that it is easily

22

repeatable. Another desire, whoever not a requirement, is regarding the time of assembly of the

spacecraft, such that it can be done quickly. Although subjective, the method used to meet this

expectation was to evaluate different assembly options such that the speed of assembly was considered.

Requirement 9.3 pertains to the purpose of the GSE. Section 3.6 will demonstrate how the author has met

this requirement with the protective enclosure.

3.3 Assembly procedures and good practice

3.3.1 GNB general assembly phases

The assembly of the GNB satellites from structural component units to full flight assembly was done

through multiple phases. The first step before conducting any assembly is to inspect every structural

component to ensure that they are within engineering specifications and tolerance. Common machine

shops mistakes should be closely looked for. Such things to take note include: proper materials, thread

size, depth and locations, corner radiuses and locations, wall thickness, opening dimensions and proper

corner filling. Following the inspection, all structures must be cleaned such that they are dirt free and no

metal particles could interfere with the assembly. The first assembly activity involved a simple fit-check

where all components were assembled to their appropriate location according to the established assembly

order, described in the following sub-Section. This fit-check has an objective of confirming the structural

design of the satellite and that all components can be properly assembled with the established tools. This

allowed a first validation of the assembly procedures such that the assembly order specified was feasible.

Figure 7 shows the fit-check assembly done for the BRITE satellite. As shown, some minor issues were

detected during the fit check including an interfering screw head and interference between the VHF shield

plate and a mounting hole.

Figure 7: BRITE fit-check assembly with some corrected anomalies

Following the fit check, a full test assembly, including wiring harness integration, was conducted. This

phase serves mainly to demonstrate the integration strategy was sound and that the full satellite can be

assembled with its integration harness in place. This acted as a second validation of the assembly

procedures. During this step, uncertain wire lengths were established and recorded such that they can be

23

replicated for future missions. For the BRITE mission, a first test assembly was done on the spare

structure to validate wiring integration strategy. This was then followed by a second full test assembly on

the flight structure to test and establish wire lengths for the panel harnesses once the solar cells were

bonded, as described in Section 3.4. As these test assemblies were done with the flight harness,

connection to the computer boards was not performed to avoid unnecessary connector cycling. Both first

and second test assemblies are shown in Figure 8. These efforts were performed by the author with the

assistance of Michael Greene, which performed most of the wire preparations including crimping and

splicing of wires.

Figure 8: First and second test assemblies

After this test assembly was completed, all components to be assembled needed be cleaned to a flight

worthy level as described in Section 3.3.2. In the clean room, a first flight-like complete assembly was

conducted while carefully following every step in the assembly procedures. This ensured that no

important steps and notes were omitted and that important observations were noted. A full safe-to-mate

check was performed on all harnesses before connecting any connectors, such that the correct pin-out for

all wires was validated. A functional test on all components will be performed during this first flight-like

assembly. Only after disassembling the satellite once more and having updated the assembly procedures

from the comments noted during the last assembly effort will the satellite be ready for full flight

assembly. Throughout all these assembly phases, rigorous documentations through the form of pictures

and commented documents were done, as will be described in Section 3.3.2. Table 4 summarizes the

GNB assembly phases used for the assembly of BRITE.

24

Table 4: GNB Assembly phases

Assembly phase Description

Structural

inspections

Ensure all structures meet engineering specifications and no sharp corners and burrs

are present.

Cleaning of

structures

Complete cleaning of structures to remove dust and particles for proper assembly as

in Steps 1 to 5 of Table 5.

Fit-check Structural assembly with appropriate tools and hardware

Test assembly Full assembly according to assembly procedures order, including harness routing

validation and wire length determination. This was done in two attempts for BRITE.

Cleaning for

flight

Complete cleaning of all components as described in Section 3.3.2

Harness fit-check Inspection of pin-out for all harnesses.

Flight-like test

assembly

Complete assembly of satellite and flight harness integration while following

assembly procedures step-by-step and documenting any divergences between actual

assembly and documented procedures. Connector cycling must be tracked.

Functional testing is also included in this step.

Flight assembly Complete full flight assembly as covered in the GNB assembly procedures document,

including all inspections, full functional testing and staking procedures.

3.3.2 Flight ready cleanliness

Although seemingly trivial, cleanliness during assembly is crucial to the proper functioning of the

satellite. This is especially true for the BRITE mission, where any dust particles could obstruct the optics

of the telescope and jeopardize the mission. Much effort was spent to ensure that all components to be

assembled were cleaned to a high standard. The flight assembly of the satellites is always performed in

SFL’s Class 10000 clean room, where the air is filtered to reduce the amount of impurities in the air.

Proper attire including an antistatic coat, shoe covers, a hair net and latex gloves is worn at all times

inside the clean room in order to reduce potential contamination and particles from the person’s body and

clothing. The general process for cleaning components includes wiping the surfaces with clean room

grade wipes and swabs moist with 99.5% pure propanol. SFL makes use of propanol-2 which leaves only

0.001% of residue after drying. Also, for mechanical components, the component is soaked in a propanol

bath in an ultrasonic cleaning machine to loosen any impurities that cling to surfaces. For printed boards

without mounted components and wiring harnesses, a propanol bath is conducted in a clean tray as

opposed to the ultrasonic machine. A complete wipe down of the component is required after any bath.

25

Also, compressed air between 20 and 40psi is blown over all components to remove any residual dust and

lint debris by the use of the labs air compressor. Before being allowed into the clean room, every

component and tool to be used must undergo a thorough visual inspection to confirm its cleanliness. A

quick assessment of the state of the component should also be performed during this inspection to look

for obvious damage or irregularities. Table 5 described the cleaning procedures which must be conducted

before an object is considered clean. When an object is clean, it must be stored inside the clean room as

soon as possible in order to maintain its cleanliness.

Table 5: Cleaning procedures

Order Process Tools

1 First blow all surfaces with filtered compressed air between 20 to 40psi in

pressure.

Filtered air

compressor

2 For mechanical components, soak the object for a period of 10 minutes in

the ultrasonic alcohol bath. This step must not be performed for electrical

components. Bathe any printed boards without mounted components and

wiring harnesses in a clean aluminum tray with propanol.

Ultrasonic

cleaning bath or

clean aluminum

tray

3 Clean all surfaces and holes such that no visible dust, lint or particles are

present using wipes and swabs with 99.5% pure propanol.

Clean-room wipes

and swabs

4 Perform a final cleaning with filtered compressed air to remove any

remaining contaminations

Filtered air

compressor

5 Perform a visual inspection. Repeat any of the other steps if still not clean.

6 Once clean, store in clean room.

Every structural component, including the trays and panels, had to be filled, thoroughly cleaned and

inspected before being accepted in the clean room. In order to protect the threads from screw insertion

cycling, stainless steel helical inserts1 are used in every thread. Although both locking and free running

Helicoils are used depending on the screw type used, prevention for loosening of the screws under

vibration is ensured by staking the head of the screws with a Room Temperature Vulcanizing (RTV)

compound. It is essential that all Helicoils be inspected after being installed such that no remaining tangs

are present, which could interfere with screw installation or become a floating particle in space which

could damage the reaction wheels, create shorts on electrical components or obstruct optical devises.

Also, every edge and hole on the structures must be inspected for sharp edges. Failure to file these edges

1 SFL makes use of Emhart HeliCoil Insert Systems for aluminum thread protection. www.emhart.com.

26

could damage the assembled harness or cause metal burrs to break off and become undesirable floating

particles like the mentioned Helicoil tangs. Figure 9 shows the process of cleaning every surface and

orifices with clean swabs and wipes moist with the propanol-2 solution. Figure 10 shows three complete

sets of structural components for the BRITE mission. As mentioned in Table 5, compressed air will be

blown over these structures a final time before they are admitted and stored in the clean room.

Figure 9: Cleaning steps for structures including hole and surface cleaning.

Figure 10: Clean structures of three complete sets for the BRITE satellite

Another cleanliness issue worthy of mention is regarding labeling of the wiring harness. Maintaining the

harness clean proved to be challenging. Dust and crimping metal debris would cling to the harness, even

after being cleaned with alcohol and compressed air. An alcohol bath was necessary to separate such

particles from the wires. Bonding labels to separate harnesses with protective tape before being put in the

alcohol bath proved to be a mistake as the glue on the tape would dissolve in the alcohol and leave glue

residue on the wires, as shown in Figure 11. It is thus required the printed Kapton labels for wire

identification be placed in the clean room once the wire is considered clean and will not be re-cleaned in a

bath.

27

Figure 11: Example of glue residue on wire

3.3.3 GNB Assembly procedures document

As mentioned, a document called the GNB assembly procedures [25] written by the author described the

full assembly procedures of the GNB for BRITE and CanX-4/-5. In the most recent version,

contributions from Stephen Mauthe allowed this document to also include procedures for the AISSat-1

payload. The document is separated in two main Sections: sub-assembly (Section 4) and final assembly

(Section 5). In the sub-assembly Section, the assembly of the different components to the two trays and

the six panels is described, as well as the magnetometer sub-assembly and the BRITE payload sub-

assembly. The assembly Section covers the step of combining the different sub-assemblies, thus the

trays, panels and payload, such that the full satellite is assembled. An appendix is also included, which

covers important information referred throughout the text; such as mounting points, harness routing

tables, computer layouts and harness protection. The assembly document made use of the Solid Edge

CAD models used by SFL for the design of the satellite.

A summary of the assembly procedures is presented in Table 6 where the sub-assembly, Section 4, is

divided into the various subsections relevant to tray, panel, magnetometer and payload assemblies.

Section 5 covers the final assembly procedures. The steps presented here only describes the general

sequence of the assembly procedures. The harness and tie-wrap assembly steps are not shown but will be

described in greater detail in Section 4.4.3. Much more detail is presented in the assembly procedures

document [25] concerning fastening hardware, verification steps, specific component installation order,

harness installation details, caution notes, tools required and more complete illustrations of the

procedures.

28

Table 6: Summary of GNB assembly procedures

Assembly Steps Illustration

Section 4.1

-Z Tray

Assemble in the following order:

• S-Band radio

• UHF radio

• Stake radio screws

• GSE support legs

• Tie-wraps (not shown)

• Radio harness partial assembly

• Radio cover

• Battery assemblies

• BCDR partial harness assembly

• GPS board

• Separation switches

• Sep. Switch partial harness ass.

• Reaction wheels (RW)

• RW partial harness assembly

• Install +/- Y Sun sensors

• Tighten and stake screws

Section 4.2

+Z Tray

Assemble in the following order:

• Tie-wraps

• VHF radio/ stake screws

• ISL radio/ stake screws

• Rate Sensor

• Radio Cover

• GPS antenna

• GSE support legs

• Imager

• Computer Board Stack

• Partial assembly of mission

specific harness

• Tighten and stake screws.

29

Section 4.3 -

Section 4.6

X and Y Panel

Assemble in the following order

(if applicable):

• Panel to GSE

• Solar cell and Sun sensor filter

verification

• Install ISS

• Install magnetorquer (+X/+Y)

• Wiring harness routing

• Mount sun sensor (X panels)

• Tighten and stake screws

Section 4.7 -

Section 4.8

+/-Z panel

Assemble in the following order

(if applicable):

• Panel to GSE

• Solar cell, S-Band and Sun

sensor filter verification

• UHF connectors with RF

terminators

• Install magnetorquer (-Z)

• Sun sensor

• VHF antenna connectors

• Wiring harness routing

• Magnetometer (+Z)

• Tighten and stake screws

Section 4.9

BRITE

payload

Do the following for the BRITE

payload:

• Mount telescope support

structure to the telescope

• Install Star-tracker mounting

bracket

• Install Star-tracker to mounting

bracket

30

Section 4.10

Magnetometer

• Insert screws from bottom

• Slide first set of spacers

• Install magnetometer board

• Slide second set of spacers

• Put on cover

• Take out screws and fasten one

by one from top

Section 5

Final assembly

• Assemble Payload and –Z tray

to +Z tray

• Complete Payload harness

• Complete BCDR harness

• Complete Reaction wheel

harness

• Complete radio harness

• Tighten and stake screws

• Assemble –Z panel and coaxial

Cables

• Assemble sun sensor harness

and –Z panel harness

• Tighten and stake screws and

connectors

• Assemble +X and –X panels

• Assemble +Z panel and coaxial

cables

• Install ISL coaxial cables

• Assemble +Y and –Y panels

• Tighten and stake screws

• Install test port cover

• Assemble UHF antennas

• Assemble VHF antenna

31

The approach taken for writing this document was to describe a generic assembly procedure where all the

possible components the GNB can handle is described in the proper sequence. The document then

specifies for which mission each component is necessary. Separate Sections are then presented for the

integration of the payload since these are mission specific. The same approach was used for the wiring

harness, where a generic harness assembly is presented wherever it is common for all missions. Harness

differences are presented where the satellites differ. Although the GNB document was initially written

for the BRITE and CanX-4/-5 missions only, the generic natures of the GNB design made it possible to

easily include AISSat-1 in the most recent version of the document, where specific AISSat-1 payload

assembly steps developed by Stephen Mauthe were included.

The method used for presenting procedures for a particular step first includes an introduction of what the

step will involve, followed by point-form directives including necessary tools and hardware. The step by

step list is then followed by a CAD illustration to better visualize the assembly process. The steps are

written such that each listed directive can be done separately and verified. Through the use of the camera

(�) symbol, the text makes frequent reminders to photograph the work. The justification used by the

author for instances when to include the camera symbol is described in Table 7. Any mission specific

steps are presented in green for BRITE, or blue for CanX-4/-5. The hardware required during every step

is presented in brackets and illustrated in the image. Figure 12 shows an example of the procedures

presentation architecture. This example refers to the assembly of both plus and minus Y panels.

Table 7: Guideline for when to take photographs during assembly

Photographing instances

• At the beginning of every sub-assembly so that a before picture of the component and structure is available

• Any part of the components being assembled which will no longer be visible once the particular assembly is complete

• After every stacking operations • After closing a tie-wrap • For verification of sun sensor filters • General pictures of the harness during harness assembly • After the completion of every sub-assembly

32

Figure 12: Procedures presentation system

3.3.4 Assembly procedures revision history

As with all SFL documents, rigorous change tracking and version control was necessary, especially since

this assembly document was also to be used by the Austrian group in the assembly of the BRITE-Austria

satellite. Table 8 describes the assembly procedures revision history including the major changes leading

to the revisions. As with all documents at SFL, the first draft of this document underwent a complete peer

review, which resulted in the first revision. Revisions 1.2 and 1.4 resulted from observations made during

various phases of the assembly, as was described in Section 3.3.1. Changes found in Revisions 1.1 and

1.2 also reflect modifications made on the board layouts or changes in the harness architecture. Revision

1.3 was done mainly because of programmatic decision to accelerate satellite assembly by partially

33

assembling the harnesses during sub-assemblies rather than wait for full assembly. This will be explained

in Section 4.5.2. Revision 1.5 includes instructions on photographing the assembly progress and to reflect

minor changes such as moving the separation switch and radio harness partial assembly to the –Z tray.

Revision 1.6 includes AISSat-1 specific steps developed by Stephen Mauthe so that the document can be

used for the AISSat-1 assembly.

Table 8: GNB Assembly Procedures document revision history

Version number Revision justification

1.0 First draft

1.1 Revision after peer review. Sections added for BRITE payload assembly, ISS, ISL,

CNAPS. Harness revision due to board layout changes. Revisions based on new

GSE design (see Section 3.6).

1.2 Revision based on structural fit-check and first assembly test observations.

Complete revision of harness due to implementation of OBC redundancy

architecture (see Section 4.5.1).

1.3 Revision to reflect change in strategy to partially assemble harnesses during sub-

assemblies and complete in full assembly (see Section 4.5.2). First version sent to

Austria for BRITE-Austria assembly.

1.4 Major revision after first flight-like test assembly. Changes to reflect use the dual

board design for CanX-4/-5 payload boards (see Section 4.5.3). Full documentation

of track changes provided.

1.5 Final revision with adjustments to UHF radio and addition of photo instructions.

Assembly of the radio harness is now partially assembled to the –Z tray (see Section

4.5.2) instead of +Z tray.

1.6 AISSat-1 specific assembly steps added. Other minor changes.

Since the Graz University of Technology is also making use of this document during the assembly of

BRITE-Austria satellite, it is necessary to provide an easy way to identify the changes made from one

version to the next. For this purpose, a copy of the document with added comments pointing to specific

changes was provided for revision 1.4, 1.5 and 1.6. An example of such change tracking is shown in

Figure 13. In addition, every SFL document requires a revision history list at the beginning of the

document, which was done for all revisions. Providing both this revision history and the extra document

copy with added modification comments proved to be very practical for update tracking by the Austrian

team, based on feedback received.

34

Figure 13: Example of commented changes in copy of revised document

3.3.5 Assembly process documentation

During the assembly process of the satellite for the flight-like test assembly as described in Section 3.3.1,

the assembly procedures document was followed point by point. The assemblers of the BRITE satellite,

which mainly consisted of the author and Karan Sarda, took note of any anomalies or additional notes that

should be included in the procedures documents. Careful tracking of the hardware used, such as screws

and washers, was compared against the hardware specified in the document. Figure 14 shows an example

of the commenting method used to specify which step was completed (�), omitted (�) or to be removed

(�). This process permitted to both record the assembly progress as well as provide improvements for

the next revision of the document.

Figure 14: Assembly tracking documentation example

Every time an assembly effort was conducted in the clean room, the progress done was recorded;

including the date, assembling members, issues encountered and reference to pictures taken. Also

recorded was the mass of every component assembled in order to verify the predicted mass of

35

components. All pictures taken were saved into the official project document folders of their respective

missions, separated into logical sub-folders based on the date and assembled components.

3.3.6 Assembly good practice

From shared experience by Cordell Grant from the AIT of CanX-2, and the author’s own experience

during the BRITE and AISSat-1 assembly, many valuable lessons were learned pertaining to assembly

good practice. These measures can be separated into three different aspects; preparation and storage,

surrounding and body awareness and safe assembly techniques. These measures are crucial for the safety

of the satellite and its efficient assembly. Table 9 lists good measures to be followed for safe and efficient

assembly. Constant reminders between the author and Karan Sarda were very useful in making sure that

proper measures were observed during the assembly sessions for BRITE and AISSat-1.

Table 9: Good assembly practice

Assembly Aspects Good measures

Clear understanding between assemblers of the planned assembly task Always put unused objects and tools out of the working area Have all hardware labeled and organized Have an organized working area Have all documents and pictures necessary available in the clean room Ensure that all documents to be used are up-to-date and accurate Store objects such that they are protected (solar cells facing down)

Preparation and

storage

Never store objects somewhere that would be prone to fall or roll Be aware of relative position of all present assemblers Constant awareness of body part resting locations, such as unoccupied hands Location of surrounding objects Location of satellite appendages such as antennas and magnetometer Always communicate intents to other assemblers Never conduct abrupt movements Never move limbs over the spacecraft unnecessarily

Surrounding and

body awareness

Track motion of moving limbs Before installing any component, do full inspection Take photographs of components before and after assembly Ensure that mechanical hardware (fasteners, washers, etc.) used is correct When fastening a component, start by loosely tightening all screws Always hold tools with both hands Never leave components unfastened when resting on structure Record all anomalies and observations

Safe assembly

technique

Make proper use of GSE as intended

36

3.3.6.1 Preparation and storage

Proper preparation for the planned assembly activity is crucial, not only for efficiency in time spent in the

clean room, but also for ensuring that the assembly will be conducted properly. All documents to be used

must be easily accessible once in the clean room and must be up-to-date. Making use of printed pictures

strategically posted on the transparent curtains enclosing the clean room proved to be very useful and time

saving. The working area must only contain tools and components that will be used in the current

assembly session. All components not used should be stored in there respective GSE such that they are

safe from their surroundings. In particular, panels should be stored in their GSE with the solar cells

facing down, so that any falling objects would not fall directly on the fragile solar cells. Any object

stored should be placed such that they cannot fall or roll. This is a particular concern for storing any

objects on grid shelves. For ease in identifying the correct screws and other hardware used, it is necessary

for them to be properly organized and labeled, since using the wrong screw could jeopardize the mission.

Figure 15 shows how the hardware sorting was done for the BRITE mission, where every screw and

spacer was sorted separately by size. Also, proper storage of components is shown, with the harness and

panels as examples.

Figure 15: Mechanical hardware sorting and proper storage examples

3.3.6.2 Surroundings and body awareness

When the assemblers are in close proximity to flight components, it is absolutely crucial that they be

aware of the location of components and each other. Constant communication between assemblers is

very important in order to be aware of each others intended movements. By not paying careful attention

of surrounding objects, it could be very easy to damage flight components, especially the sensitive

exposed solar cells or delicate antennas. When resting your hands on the working table, the space used

must be free of objects. Whenever an assembler must reach for a tool or object he must be aware of the

path of his arm such that it does not pass over or in proximity to the satellite unnecessarily. Abrupt

movements should be avoided to reduce the risk of hitting the spacecraft. Good awareness of

37

surroundings and mastering safe body movements comes from practice and constant reminders between

assemblers through constructive observations.

3.3.6.3 Safe assembly techniques

During the assembly, certain measures must become habitual so that safe and repeatable assembly can be

performed. Before assembling a component, a complete inspection of the component and the structure to

which it be assembled to must be performed. Things to observe include cleanliness, edge sharpness and

overall state of the component. The component should be photographed before and after being

assembled. This will permit future reference to the time at which specific anomalies became apparent.

Any anomalies or observations must be recorded in the assembly logs for future reference. When

assembling the component, the tool used, such as a screwdriver or a pair of tweezers, should be held with

both hands to reduce the risk of dropping it on the spacecraft. When assembling a component, such as a

computer board for example, the screws holding the computer board should first be loosely fastened so

that it is not subjected to unnecessary stress from assembly misalignments. Only when all screws are in

place should they be tightened. Never leave any components resting on a structure without being

fastened. As an example, the magnetometer boom should not be resting on the panel if it is not fastened.

Inadvertently knocking over the magnetometer boom could damage the solar cells on that panel. A final

aspect worthy of mention is the proper use of the GSE. The GSE used as an assembly aid should always

be used as it was intended. If it is lacking for a particular application, its design should be revised. This

approach was used for the design of the GSE, as will be presented in Section 3.6. Although these are

simple measures, they should always be performed in order for the satellite to be properly and safely

assembled and for the source of any anomalies to be easily determined.

3.4 Solar cell and S-Band patch antenna bonding pro cedures

3.4.1 Use of bonding compound

The bonding compound selected for the GNB solar cells is the NuSil Technologies CV-25662, which

proved to be effective during the CanX-2 solar cell bonding. This Controlled Volatility (CV) silicon RTV

is a two-part (base, curing agent) rubber compound specifically designed to bond solar cells to metal

surfaces in a space, high vacuum, environment. The Nusil SV-1213 primer is purposely formulated to

improve adhesion of the NuSil RTV compounds to metals. The general process for proper curing of these 2 http://www.nusil.com/library/products/CV-2566P.pdf 3 http://www.nusil.com/library/products/SP-120_SP-121P.pdf

38

products includes applying a thin layer of primer to the clean and dry metal substrate, let dry for 30

minutes, mix the RTV base with its catalyst at a weight ratio of 100:0.5, bond the solar cell, and let cure.

At SFL, curing at room temperature for 24 hours with pressure on the solar cells from small weights (1 to

1.5kg), followed by 48 hours at 80°C without weights, proved to be an effective curing scheme.

3.4.2 Solar cell bonding procedures summary

Based on the experience gained during the CanX-2 mission, full procedures were developed for the

bonding of the solar cell coupons to the panels. A key difference between CanX-2 solar cells is that GNB

solar cells came pre-assembled to an aluminum substrate. This simplified the manipulation of the solar

cells for the GNB. A summary of the full procedures will be presented in this section. Table 10 shows

the solar cell bonding abbreviated procedures in steps, with images to aid in understanding. If engaging

in solar cell bonding for a GNB mission, the full document should be followed [26].

Table 10: GNB solar cell bonding procedures summary

Steps Description Illustration

Working area

preparation

Find wide open clean area

where the solar cell bonding

will be done.

Preparation • Clean all panels

• Cover all holes with a small

square piece of tape

• Make use of mock coupons

and rulers to determine area

used by solar cells

• Place two layers of tape to

delineate the area to be used

by the solar cells and set

proper RTV thickness

39

Work area

organization

Place panels in central location

with solar cells to be bonded on

one side and all tools on the

other. Make drawings of panel

assembly easily accessible.

Solar cell

preparation

• Unlace cell wires

• Place solar cells in expected

position

• Perform fit check with panel

• Take photos of solar cells

with number tag and record.

Primer

application

• Apply primmer to both

panel open surfaces and

back of solar cell coupons

with clean-room wipe

• Let dry for 30min

Prepare RTV

and degas

• Mix the RTV with its

hardener according to

specified ratio of 1:200

• Degas RTV in bell jar

Spread RTV

on panels and

place solar

cells coupons

• With a plastic clean spatula,

apply RTV in a thin layer,

as thick as tape delineation

on all panels.

• Place all coupons in their

approximate location

• With sharp edge, precisely

40

adjust location of coupons

Apply

cheesecloth

and sponge

• Cover solar cells and border

RTV with vacuum bagging

cheese cloth

• Place two sided sponge with

soft side down

Place weights

on sponge

• Place 1 to 1.5kg on each

coupon such that weight is

evenly distributed

• Let cure at room temp. for

24h

Remove

cheesecloth

• Remove weights, sponge

and cheesecloth

• Inspect coupon placement

Bake panels

in

temperature

chamber

Place panels with solar panels

in temperature chamber at 80°C

for 48h.

Remove tape

and clean

• Inspect state of solar cells

• Remove all tape

• Remove RTV residues with

sharp knife and propanol

moist wipes.

41

Install in GSE

and place

wires

• Fasten the panel to its panel

GSE

• Pass wires through holes

The most important aspects to keep in mind during the bonding of the solar cells are; to keep the working

area, panels and solar cells organized and clean, properly document the progress, not apply excessive

amount of RTV and ensure proper placement of solar cells before covering them with the cheesecloth.

The amount of RTV applied should be no thicker than the two layers of Kapton tape which bounds the

RTV application area. Special care must be applied when placing the weights such that they do not move

the solar cells and are correctly position such that they will not fall off the sponge. Care must be taken

not to apply excessive force on the fragile solar cells. After the first room temperature curing, the RTV

that was remaining in the mixing dish should be inspected to confirm proper curing, thus validating

proper mixing with hardener during the RTV preparation.

3.4.3 S-Band patch antenna bonding

The bonding for the S-Band antenna patches onto the panel is essentially identical to the Solar cell

bonding process. The only particularity to consider is the extra step of temporarily covering the central

opening in the panel where the antenna connector passes through. Figure 16 shows this extra step. As

described in procedures document [26], it was observed that time could be saved if the bonding of the S-

Band patch antennas be done during the same session as during the solar cell bonding. This will also

reduce the risk of damaging the already bonded solar cells during the bonding of the S-Band antenna.

Figure 16: S-Band patch antenna bonding

42

3.4.4 Use of excessive RTV for AISSat-1

Since the results were favorable and no issues occurred during the solar cell bonding for the BRITE

mission, these procedures were demonstrated to be sound. These same procedures developed by the

author were later used by Stephen Mauthe during the bonding of the AISSat-1 solar cells. Some solar

cells had to be removed and re-bonded since they had shifted from the desired position. This issue was

agreed to be caused by an excessive amount of RTV used during its application on the panels. Hence, the

document has been revised to include a note cautioning that it is better to use too little RTV and add as

required, rather than applying too much and attempting to remove the excess as was done on the first

AISSat-1 solar cell bonding effort. If the Kapton tapes delimiting the application area are completely

covered with RTV, the amount of RTV is excessive.

3.4.5 Solar cell inspection and functional testing

Full inspection of all the solar cells under microscope must be done before the solar cell bonding and after

thermal chamber curing in order to identify any micro-cracks or solar cell defects. The author assisted

Mihail Barbu for the inspection of the BRITE solar cells, where any small noticeable anomalies where

recorded. A simple functional test of measuring the solar cell voltage output under illumination was

performed and recorded by the author as shown in Figure 17. The solar cells used on the GNB missions

are 26.8% Improved Triple Junction solar cells from Sprectrolab4 with an Open Circuit Voltage (Vop) of

2.565V. The voltage output from the high resistance multimeter should thus read somewhere close to 5V

as each coupon is composed of two solar cells in series. More importantly, it should be consistent

between solar cell pairs on the different panels. This was the case for the BRITE solar cells. This test

gives a good indication that all solar cells behave as they should.

Figure 17: Solar cell short functional test

4 http://www.spectrolab.com/DataSheets/TNJCell/tnj.pdf

43

3.5 Thermal tape application

In order to achieve the desired thermal characteristics of the spacecraft in orbit, the thermal engineer for

the GNB, Karan Sarda, made use of gold covered Kapton5 tape. In order to apply this tape to the panels,

the author, along with Karan Sarda, developed a simple technique for placing and applying the tape [27].

The first step is to measure and cut the tape to the required dimensions. Care must be taken not to scratch

the surface as this could change the thermal characteristics of the surface. With the use of a plastic

spatula, the thermal tape was applied, taking care that no bends or air bubbles would get trapped between

the tape and the panel. This trapped air could de-bond the thermal tape in a vacuum environment.

Although more than one attempt was necessary at many instances, this technique proved to be simple and

straightforward. With the use of a sharp knife, tape covering every opening and hole was removed, taking

care not to cut any nearby wires or damage the edge of the opening. This was particularly important for

the BRITE telescope opening, where damage to the edge of the orifice could interfere with proper optics.

Figure 18 shows some of the steps involved in the thermal tape application, and Figure 19 shows an

example of the resulting covered panel and magnetometer boom.

Figure 18: Thermal tape application techniques

Figure 19: Panel and magnetometer boom covered with gold thermal tape

5 Kapton is a Dupont registered material: http://www.dupont.com/kapton

44

3.6 GNB ground support equipment

As briefly explained in Section 2.5, the Ground Support Equipment (GSE) is used to both facilitate the

assembly of the spacecraft and protect it once it is assembled. This GSE is composed of tray and panel

support jigs, Lexan protective panels, a magnetometer boom protective cover, tray support legs and a

“Lunch Box” support jig. What is referred to as the Lunch Box is the protective enclosure cube

composed of tray support jigs with attached Lexan panels which encapsulates and protects the assembled

satellite. A first iteration of the Lunch box was conceptualized by Benoit Larouche [7]. The author

revised this early design and produced the final version which was used for the GNB missions. The early

concepts, final design, and its manufacturing will be covered in this Section.

3.6.1 GSE requirements

In addition to the generic systems requirement SYS 9.3 in Table 3 relevant to the GSE, a set of

requirements was developed for the design of the protective enclosure. Table 11 lists such requirements.

Table 11: GSE Requirements

GSE requirements

GSE 1 The GSE shall enable access to both sides of the GNB trays during their sub-assembly such

that no mounted component comes in contact with anything but the GSE.

GSE 2 The GSE shall enable access to both sides of each panel such that no panel components

come in contact with anything but the GSE.

GSE 3 The protective enclosure, or “Lunch Box”, shall completely enclose the fully assembled

satellite such that it is protected from its surroundings (with the exception of the antennas)

GSE 4 The GSE shall enable access to any side of the enclosed satellite without the spacecraft’s

disassembly.

GSE 5 Each panel of the GSE shall be easily dissembled without the need of disassembling any

other panels.

GSE 6 Openings shall be provided to account for any pre-deployed appendages (ie. magnetometer

and antennas)

GSE 7 The materials of the GSE shall support temperatures ranging from 20 to 50°C

GSE 8 The materials of the GSE shall not react with the cleaning agent or anything in the satellite

45

3.6.2 Early Design

The early designs of the GNB were much more elaborate, thus more machining intensive versions of the

final version used. Also, as seen in Figure 20, the tray support legs were initially attached to the tray

support jig rather than directly to the trays themselves. From discussions, it was preferable to handle the

structure directly rather than being supported by the GSE when it is being rotated to reduce risk to the

satellite. The support jig for the trays would have been made of a single piece, which would have

required considerably more machining. For that reason, the support legs are now attached directly to the

tray. In terms of panel support jigs, the initial design suggested making use of a single support jig which

would have been able to support any panels. Although this would have limited the support jigs to a single

version, it was not cost effective and required significantly more machining than the final product which

includes a dedicated panel jig for each panel set. During these early designs, the protective enclosure (or

“Lunch box”) had not yet been implemented; hence this concept would have required some further

revisions.

Figure 20: Early designs of the GSE

3.6.3 Final Design

The final design of the GSE is made from simple machined Delrin bars for the support jigs and Lexan

plastic for the Lunch box panels. By keeping machining to a minimum while satisfying requirements, it

was possible to manufacture the GSE entirely at the UTIAS machine shop. In order to satisfy

requirement GSE 1, tray support legs and rails assembled to a Lexan panel was produced as shown in

Figure 21. The length of the support legs is determined by the clearance height required by the UHF and

S-band connectors when these radios are assembled. They are shaped such that they may not rotate when

46

assembled. The tray support rail constrains the tray in two directions and allows the tray to rest in the

support jig with an assembled Z panel and RF terminators in place.

Figure 21: GSE Tray support rails and legs

In order to satisfy GSE 2, simple Delrin blocks with assembled aluminum spacers and screws were used

to hold the different panels such that they may rest in either normal directions. A total of 4 different

support jigs are used for the X, Y, +Z and –Z panels as shown in Figure 22. It also shows the +Z and –X

panels assembled to their respective support jig.

Figure 22: GSE panel support jigs

The protective enclosure is composed of two assemblies of tray support rails with the addition of four

additional Lexan panels to fully enclose the satellite. A magnetometer boom protective cover is

assembled with the use of screws to the +Z tray support jig which include Helicoils for thread protection.

It can be disassembled from the support jig easily. Finally, a lunch box raising platform was

manufactured in order to satisfy GSE 4, where any sides of the satellite enclosed in the lunch box may be

accessed in the event that the UHF antennas were assembled. Figure 23 shows the protective enclosure

with the raising platform and photos of the GSE about the BRITE mass dummy on the left and BRITE on

the right.

47

Figure 23: GSE Protective Enclosure

A later variant of the support legs were manufactured in order to adapt to a new UHF version which made

use of longer connectors, thereby increasing the height by which the tray needed to be raised. This design

also made use of a cross brace between the legs in order to improve the support rigidity. It was also found

to be better suited for the partial assembly of the harness to the trays as they did not interfere with the

assembly of some of the tie-wraps as was the case for the first support legs version.

Figure 24: New GSE support legs variant

3.6.4 Materials used

The materials selected for the manufacturing of the GSE was done with protection of the satellite in mind.

It was demonstrated from the GSE used for CanX-2 that Delrin6 was appropriate for protecting the

6 Delrin is a brand name of Dupont.

http://www2.dupont.com/Plastics/en_US/Products/Delrin/Delrin.html

48

aluminum structures of the satellite. The choice of Lexan7 for the panels of the GSE was mainly based on

the need of transparency and its ease of machining, as well as its demonstrated suitability from the

CanX-2 GSE. Lexan Polycarbonate tubing was also used for the magnetometer cover, mainly for its

availability and suitability to the task. All of these materials satisfy requirements GSE 7 and 8 and are

described in Table 12.

Table 12: GSE materials

Material Used for Characteristics

Delrin • Tray Support rails • Tray Support legs • Panel Support jigs

• Operating temperature: 0°C to +82°C • Good resistance to alcohol • Smooth finish

LEXAN

Polycarbonate

• Protective enclosure panels

• Magnetometer cover

• Operating temperature of -135°C to +115°C • Good resistance to alcohol • Clear • High impact strength

3.6.5 GSE Manufacturing

As mentioned, all manufacturing of the GSE for BRITE and AISSat-1 was done by the author at the

UTIAS machine shop or with SFL equipment. Many lessons were learned from this experience for both

good design and machining techniques. Table 13 lists recommendations from such experience. A good

design always considers how the part will be manufactured. In terms of machining, this includes a good

understanding of the machine shop’s limitations and the tools available. For good machining practice, it

is essential that a clear plan on how to proceed be established before attempting machining. This includes

the setup plan and reference system to be used. Improper preparation will typically lead to more time

spent correcting setup issues and re-establishing references, which could also lead to accuracy problems.

7 Lexan is registered trademark for SABIC Innovative Plastics brand

http://www.sabic-ip.com/gep/Plastics/en/ProductsAndServices/ProductLine/lexan.html.

49

Table 13: Manufacturing recommendations

Aspects Good measures

Good design for

manufacturing

• Make use of readily available commercial material stock sizes • Design such that a minimum of machining operations is required • Whenever possible, keep recurring dimensions, like radii and holes, identical

such that a minimum of machining tools are necessary. • Design with exact dimensions with the unit type used by the machine in mind • In technical drawings, keep dimensions clear and organized such that every

dimension is given only once. There should be no ambiguity in the drawings for drawing directions and dimensions.

• Consider if mirrored and generic dimensions should be used. In many instances, machining time can be significantly reduced if dimensions are repeated, even for different parts. Care must be taken however if this generic dimensioning method can lead to assembly confusion and errors.

• Always ensure that every machining operation necessary for the part manufacturing is feasible. Take note of the degrees of freedom of the machining equipment available.

• Keep the machining directions required to a minimum to minimize the setup time required and improve relative dimension precision.

• Specify tolerances as required, without exaggerations for dimensions which may tolerate less stringent machining accuracy.

Good

Manufacturing

techniques

• Always make sure you have all the tools and setup jigs required before initiating machining.

• Make sure you understand the machine you use and all its functions. • Decide before attempting machining on the machining sequence with the

required setup jigs to be used. • Start with the most important dimensions whenever possible. • Make sure the machining operation planned can be entirely performed with the

setup jig in place without the need to reposition the part half way through the operation.

• Make use of stopping blocks for repeated dimensions whenever possible. • Always make use of a common absolute reference point or edge for all

operations rather than making use of changing reference edges. This will ensure that dimensioning error is consistent rather than compounding.

• Ensure that the edges of the part to be machined are orthogonal with the machine axis.

• Ensure that the machine is clean and the tools used are in good machining condition.

• Always make rough cuts before the final cut.

50

4.0 GNB INTEGRATION

As was mentioned in Section 3.1, integration involves interconnecting the different components in order

to make the satellite functional. This involves the wiring harness design, including its mounting points,

routing and assembly to the GNB structures. The integration requirements will first be presented,

followed by lessons learned from integration of the CanX-2 satellite, the integration strategy for the GNB,

computer board layout, RF harness design and critical design changes which affected the harness design.

Finally, consideration of an alternative integration strategy for future missions will be covered.

4.1 Integration Requirements

The author developed the GNB wiring harness requirements document [28] from which was derived the

integration requirements presented here. A revision of the document was done by Michael Greene with

the addition of splicing specific requirements. The requirements in this document mainly originate from a

technical standard document [29] developed by the National Aeronautics and Space Administration

(NASA) in the United States. Since not all requirements presented by NASA were applicable, only the

relevant requirements were included. From experience on the CanX-2 mission by Cordell Grant, personal

experience of the assembly of the GNB missions, and standard SFL practice, these requirements were

agreed to be sound and relevant. Table 14 presents these requirements with sources and compliances.

Table 14: Integration requirements [28]

Physical requirements

Requirements:

• The harness shall be composed of materials that exhibit a total mass loss of no more than 0.1% of the

component’s initial mass, and that contain no more than 0.1% collected volatile condensable material.

• The total mass of the harness shall be recorded.

Source:

• From GNB system requirements [24] and SFL standard practice in order to avoid outgassing issues.

Recording of mass is done to update the mass budgets for the GNB and have more accurate mass

estimation.

Compliance:

• All materials composing the wiring harness (including heat shrink, tie-wraps, wires and connectors)

are compliant with the outgassing requirement.

51

• Total mass of harness was measured to be 185g, as described in Section 3.3.5.

Assembly requirements

Requirements:

• The harness should permit easy assembly and disassembly of the connected components.

• The harness shall be assembled with simple, common lab tools.

• All connectors shall be stacked.

• The harness should be designed such that connector cycling is minimized below product

specifications.

• Connector cycling shall be recorded such that a specific connector has cycled less than 37cycles.

Source:

• Derived from Systems requirement and standard SFL practice.

• Manufacturer specify maximum of 50cycles for DF-11. With 75% derating, maximum is 37.

Compliance:

• The ease of assembly and disassembly is verified through assembly testing and refinement of the

assembly procedures.

• All cycling of connectors was recorded and compared with the DF-11 cycling limit of 37 cycles. The

vibration testing will ultimately verify the robustness of the connections.

Mechanical requirements

Requirements:

• No wires shall obstruct the field of view of optical devices.

• Wiring of functional redundant components (eg. BCDR) should not be routed in the same bundle or

through the same connector.

• No object shall come in the vicinity of parts that are required to move (eg. reaction wheels).

• Wiring tie-down methods that can potentially pinch wires should not be used.

• The harness shall include protection at the mounting points for stress relief and Teflon cold-flow

avoidance and electrical insulation.

• The tie-wraps used for flight shall have tangs that lock securely into the “ribbed” portion of the straps

• Harness protection shall be added in areas where sharp or rough edges are present.

• The minimum bend radius for a wire bundle should not be less than 3x the Outer-diameter (OD).

• The minimum bend radius for bundles containing coaxial cables should not be less than 6x OD.

• Straps used should be tightened such that they do not slide but not too tight as to cause noticeable

indentation in the harness.

• The harness shall be bundled and strapped with appropriate protection and tie-wraps before and after

52

every breakout for splicing.

Source:

• Derived from Systems requirement, assembly experience and standard SFL practice.

• Redundancy consideration, wire protection, tie-wrap use, protection against sharp edge, strain relief

and bend radius requirements are derived from NASA harness standard document [29].

Compliance:

• In general, proper wire protection is accomplished with the appropriate use of Kapton tape and Heat

shrink sleeves and an appropriate mounting strategy to avoid wires being stressed or taught.

• Careful inspection of wire bends ensures that its bend radius is not excessively small. The routing

strategy also considers bend radius, particularly concerning coaxial cables.

Electromagnetic Interference (EMI) requirements

Requirements:

• No wires shall form a complete loop.

• Every power wire and RS485wires shall be twisted with its corresponding return wire.

Source:

• SFL practice in order to reduce magnetic flux in the circuit and system noise (see Section 4.1.1).

Compliance:

• Appropriate routing of the wires ensures that no complete loop is produced which would produce an

undesired magnetic field and resulting magnetic disturbances.

• Twisted wires with their return wires are ensured by inspection of the harness. This is essential to

prevent undesired system noise from ambient electromagnetic generated currents.

Communication System requirements

Requirements:

• The phase match between the coaxial cables for the same radio shall respect a tolerance of ±10%.

• An RF terminator shall be used for the radio coaxial cable terminations whenever the antenna is not in

place.

Source:

• Proper signal phase is required for signal processing. Since time of signal propagation is proportional

to cable length, the 10% phase matching requirement translates into a ±10% cable length tolerance.

• RF terminators are used to protect radios for improper impedance after the removal of the antenna.

Compliance:

• Phase matching is ensured by making coaxial cables for the same radio of equal length and type, and

verifications with RF measurements of phase matching.

53

Electrical Loss requirements

Requirements:

• The voltage drop across the wiring harness shall be low enough that the voltage reaching every

component is in the component’s voltage operating range.

• The current passing through every wire shall be lower than the maximum current limits imposed by

the derated ampacity value of the DF-11 connectors/crimps.

Source:

• The voltage drop limitation is required for proper operation of components.

• Derating limits are specified by the ECSS Space Product Assurance document [30]. Current limits are

dependent on the wire gage and number of wires in the bundle.

Compliance:

• Electrical losses are first validated by FlatSat testing where the fully integrated satellite ensures that

voltage and current losses are appropriate. Functional testing on the flight assembly with flight

harness ensures compliance.

Validation requirements

Requirements:

• All harness shall be tested for point-to-point electrical continuity by someone other than the person

who conducted its assembly such that the harness pin-out is identical to the electrical inter-connection

diagram. This is referred to as a “Safe-To-Mate” test.

• Every wire shall be inspected before being assembled to ensure that no wires have exposed copper and

that the wire insulation intact

• After harness integration, every wire and connector shall be inspected to ensure that it is not under

stress such that the harness may be moved by several mm between tie-points without stretching the

wire.

• Every connector shall be inspected to ensure that it is properly mated, staked.

Source:

• Safe-to-mate check is derived from NASA harness standard document [29] and standard SFL practice.

• Validation of the integrity of wires and connectors is done by visual inspection and point-to-point

conduction with the use of a multimeter.

Compliance:

• All safe-to-mate tests at SFL are done through peer review and recorded in assembly logs.

• The assembly procedures of the GNB include inspection steps to remind the assembler.

• All inspection steps are recorded in the assembly logs.

54

4.1.1 EMI reduction by twisted pairs

In order to reduce the Electromagnetic Interference (EMI) in the system, it is required to twist pairs of

conductors with opposite currents directions. This is the case for the power/return wire pair for power

transmission and the transmit/receive pair used for the RS485 serial protocol, where a full duplex system

is implemented. The reason for this is twofold. First, the magnetic flux induced in a circuit loop from an

ambient magnetic field is proportional to the area of the current loop. A changing flux will generate an

undesired voltage, thereby increasing the noise in the system. By twisting the wires, we are effectively

reducing this area. Also, twisting the wires will lead to areas of alternating signs, which would partially

cancel the electromagnetic field generated between adjacent opposing loops [31][32]. Twisting the wires

will also reduce the capacitive coupling between conductors. A mutual capacitance is generated by

spurious electrical coupling between conductors in electrical circuits. Since capacitance is created by

differences in potential, nearby conductors will create this mutual capacitance. The mutual capacitance

(C) generated by two parallel circular cross-Section wires at a distance (ds) of each other and at a height

(h) above a ground plane is well understood [31]. It was shown [32] that C will be reduced with a smaller

h and larger ds. Similarly, magnetic-field coupled noise, or mutual inductance coupling, can also be

reduced by running conductors close to the ground plane, also achieved by twisting.

4.2 CanX-2 Integration Lessons

Following the successful completing of the CanX-2 mission, many recommendations were suggested by

Cordell Grant, the CanX project manager which was responsible for the AIT of CanX-2. These

recommendations pertained to better integration practice and design strategies. Table 15 shows the

relevant recommendations found in Cordell Grant’s thesis [2], which were followed by the author in the

development of the integration strategy for the GNB missions.

Table 15: AIT recommendations from CanX-2 development

Number Recommendation from CanX-2 AIT Actions taken for the GNB AIT 1 Preliminary design of the wiring harness

and selection of tie-down points should be performed prior to the structure being manufactured.

All tie-down points and wiring harness routing design was completed by the author before the structure was manufactured.

2 Margin should be applied to the areas designated for solar cells to account for wiring and a plan should be in place for such wiring prior to the structure being manufactured.

All wiring harness routing on panels, including location of wire through holes and solar cell locations were verified and established by the author before any structures were manufactured.

55

3 Splitting the main wiring harness into two or more pieces should be considered to make construction and integration of the harness easier. The benefits of splitting the harness should be weighed against the added risk associated with having more connectors.

The wiring harness for the GNB architecture is composed of a specific independent harness for each component and follows a specific routing defined by a tie-point sequence path. The interconnection architecture was revised with contributions from the author for improved ease of integration.

4 The AIT engineer must be consulted regarding connector location and pins assignment.

The connector layout for the Power Board and OBCs was decided from discussions between the author and the computer and power hardware designers. [see Section 4.4.6]

5 The structural engineer should be consulted regarding connector positions on all electronics boards.

Constant communication between the author and the computer and structural engineers ensured that no interference would occur.

4.3 Integration Architecture

The integration architecture for the GNB is a decentralized system where power is distributed to every

component through the power board and data is handled by either On-Board Computers (OBC). Figure

25 shows how data and power is distributed between computers and components. Although each OBC,

the House Keeping Computer (HKC) and the Attitude Determination and Control Computer (ADCC),

have dedicated functions, the integration architecture includes the ability to interchange their roles thus

adding “cold” redundancy to the system. This feature is evident by the presence of dashed data lines

going to the other OBC for the reaction wheels and radios, representing existing, but unused connections.

This is accomplished with either dedicated wires or splices. The accommodation of this redundant

approach by the harness will be presented in Section 4.5.1. Components in blue are part of the ADCS

subsystem, in purple for the communications subsystem, red for power, yellow for computers and green

for payloads. Lines in red represent power links, blue lines indicate telemetry links and green is for both

power and data. As can be seen in the diagram, the GPS is nominally connected to the payload computer,

the power board and the AIS sensor in the case of AISSat-1. Connections from the GPS receiver to both

OBCs also exist but nominally will not be used. All antennas make use of coaxial cables for signal

transmissions to their respective receiver or transmitter. Rather than a direct connection to the OBCs,

some components such as the imager and the magnetometer will be connected to the Power Board and

then be redirected to the OBCs through the power board-OBC links.

56

Figure 25: GNB Integration Architecture

4.4 Integration design

The author was responsible for designing the wiring harness for the GNB which would complete the

integration of the satellite. As was mentioned in Section 4.2, many lessons which were learned during the

CanX-2 integration process by Cordell Grant were followed in order to design the harness for the GNB.

This Section will first present the wire and connector types used for this harness, which were based on the

CanX-2 harness and pre-established before the design of the harness. Different harness approaches were

developed in parallel with the design of the computer board for which the author provided ideal connector

locations for ease of integration. The validation of the harness design was verified through the solid

models, expected bundle size estimations, and physical test assemblies as described in the assembly

Section 3.3.1. The routing of the coaxial cables was also developed such that communication and bend

radius requirements were satisfied. For integration repeatability, a documentation system was developed

and used in the GNB assembly procedures [25] as will be discussed in this Section.

57

4.4.1 Wiring harness components and materials

The wiring harness involves several different components including: Teflon insulated copper wires,

coaxial cables, DF11 connectors, SMA connectors, MCX connectors, heat shrink sleeves, and tie-wraps.

A very important component of the wiring harness is the wires themselves, which acts as a medium for

data and power distribution in the satellite. For power, either AWG22 or AWG24 American Wire Gages

(AWG) are used. For telemetry, AWG28 is used. These wires8 are rated to temperature ranges from -

65°C to 200°C, are chemically inert, unaffected by solvents, and are rated to 600V. In terms of

connectors, the Hirose DF-11 connector was used for its low profile and availability. It meets the

outgassing and thermal requirements listed in Section 4.1. It is rated to a maximum of 50 mating cycles

and meets the lower then 30mili Ohms resistance specification. According to the SFL quality assurance

plan [33] which follows the MIL-HDBK-1547 military standard, the derating of electronic components is

75%. With this derating factor, SFL limits the cycling of the DF-11 connector to 75% of 50 cycles, thus

37cycles as mentioned in the requirements. For coaxial connectors, either the SMA threaded connectors

or the MCX snap connectors are used, depending on the space available. The SMA connector and coaxial

cables are also used as part of the BRITE and AISSat-1 payload harness. The coaxial cables used are of

the RG316 type, which follows the MIL-SPEC-C-17 specifications. The tie-wraps used are Tefzel9 cable

ties which are rated from -60°C to 170°C and have high chemical resistance.

4.4.2 Harness routing and bridging strategy

In order to satisfy requirements presented in Section 4.1, in particular the requirements pertaining to

efficiency and ease of assembly, the routing was designed such that the most direct route was chosen

between computer board connectors and components. With this approach, the wires lengths were

minimized, which resulted in less harness mass and less power loss through the harness. Also, keeping

wires short reduces the EMI concerns related to the harness as was described in Section 4.1.1. Different

routing concepts were considered such as having all harness bridging from one tray to the other at a

common location. This proved to be impossible given the space available for the harness between the

trays and panels compared to the volume required for the harness. This also opposed the desire of

minimizing wire lengths. Nevertheless, iterative routing design resulted in most harnesses routed from

the –Z tray to the +Z tray on either the +Y or +X face, with the exception of the BCDR harnesses on the

–Y face. This proved to be the most appropriate bridging strategy in order to properly distribute the

8 SFL makes use of Alpha Wire: www.alphawire.com 9 Teflzel is a registered trademark of DuPont. www.Dupont.com

58

harness about the satellite and permit the disassembly of the payload without the need to remove most of

the harness. Since the payload openings for both BRITE and CanX-4/-5, such as the telescope aperture

and the CNAPS valves, were on the –X face, no wires would bridge from one tray to the other on this

face. The assembly order of the harness was established in parallel with the assembly order of

components. Through iterations and test assemblies, the specific tray to which the harness should be

assembled first was established. For fixing the harness to the structure, the location of mounting points

were carefully determined while considering the ease of access of the tie-wraps and ease of their

replacement. For most cases, components can be removed without the need to break any tie-wraps

holding its associated harness to the trays. The only exception is the payload, which occupies the central

volume between trays. Because most harnesses bridge across both trays and are fixed to both trays,

disassembling the payload does require some disassembly of the harness, namely the BCDR harness. The

alternative integration concept presented in Section 4.6 can be a potential solution to avoid this

inconvenience. Figure 26 shows this bridging concept of the harness on the partially assembled BRITE

spacecraft during the second test assembly. Only the BCDR harness cannot be seen as it is on the

opposite face.

Figure 26: Photograph showing the bridging strategy used for the GNB wiring harness

59

4.4.3 Mounting strategy

The mounting strategy used for the GNB was inspired by the successful mounting strategy used for the

CanX-2 integration. It consists of wrapping the harness bundle with a heat shrinking sleeve and fixing it

to the structure with a tie-wrap which passes through two spaced holes. As will be described in Section

4.4.4, the spacing of these holes for each tie-wrap depends on the amount of wires in the bundle and the

room available for that bundle. The direction of the tie-wrap buckle when closed also depends on the

space available and the ease of access for the tie-wrap closing tool used. This mounting strategy aims at

answering the requirements of properly stress relieving the connectors, preventing motion of the harness

during vibration and protecting the harness against wear. Also of importance is the consideration of

Teflon cold flow. Cold flow is a well known phenomenon [29] where Teflon wire insulation can flow if

pressed and leave the copper conductor exposed. To mitigate this risk, heat shrink is used at every tie-

point such that the pressure by the tie-wrap on the wires is better distributed and the wires are not directly

pressed against the metal structure. Also, careful inspection of every tie-point is done to ensure that the

tie-wraps are not excessively tightened about the harness as to result in noticeable wire deformation.

Kapton tape is also used to secure the heat shrink sleeves to the wires and cover sharp edges to protect the

harness. As a general strategy, the variety of harness passing through the same tie point was kept to a

minimum in order to be able to close the tie-wrap and secure the harnesses as soon as possible.

4.4.4 Harness size verification

A physical constraint for the harness design was the physical space available for the harness bundles and

mounting points. The first validation was done during the harness design where the spacing between the

tray and panels was determined from the CAD models and compared to the expected bundle sizes as

shown in Figure 27.

Figure 27: Maximum bundle height determination from CAD model

60

A full spread sheet was produced for every tie-point to determine the required tie-wrap hole spacing based

on the empirical equation [1] developed by the author. This equation was developed as a guide in

selecting the amount of wires that should be included in a specific bundle held by a tie-wrap based on the

spacing of the holes in which the tie-wrap passes through and the maximum height available based on the

space available. If the theoretical hole spacing S (in mm) was wider than the actual spacing available for

the tie-wrap in question, than the amount of wires passing through this tie-wrap was reduced.

CN

DN

N

DN

N

DN

H

HS

refC

CC

refT

TT

refp

pP

act

ref +

×+×+×

=

,,,

2 [1]

In this empirical equation, S represents the theoretical hole spacing (in mm). Href represents a reference

bundle height of 6mm based on an average value of the space available between the tray walls and the

panels, whereas Hact is the actual bundle height available for the specific bundle considered. Np, NT and

Nc is the amount of wires of different types, representing power, telemetry and coaxial cables

respectively. Dp, DT and DC are the respective diameters of the wire type (in mm), which for the power,

telemetry and coaxial cables where 1.5mm, 1.0mm and 2.5mm respectively. Np,ref, NT,ref and Nc,ref are the

reference number of wires of the specific type which can be stacked on top of each other to reach the

reference height Href, based on their respective wire diameter. These values were 4, 6 and 2 for the

power, telemetry and coaxial cable respectively. A factor of two is added for the spacing occupied by the

power wires as they are twisted, which effectively doubles the width required by these wires. The

constant C is a correction factor to take into account the rounding affect on the bundle due to the tie-wrap

pressure and heat shrink wrapping. A value of 1mm for C was selected from iteration to best approximate

the practical size from measurements. Figure 28 demonstrates this bundle size concept with an example.

Figure 28: Bundle size approximation representation

61

The actual bundle sizes were verified and compared to the expected results during the first test assembly,

which validated the estimation process used. Although the space constraint for the harness was

challenging, these bundle size estimates permitted an appropriate distribution of harnesses across the

many tie-points while permitting the full assembly of the satellites. From the test assembly, minor

adjustments were done in terms of the mounting point selection of a few wires and the buckle direction of

some tie-wraps. Full inspection was done during the assembly of every panel in order to ensure that they

did not interfere with tie-wraps or wires, and that no wires were being pinched or compressed. Figure 29

shows this type of inspection.

Figure 29: Harness inspection during panel assembly

4.4.5 Coaxial cable routing

In order to satisfy the requirements for phase matching, the coaxial cables had to be designed such that

they are of equal lengths. Also of importance is the order of cables connecting to the UHF radio. Since

the UHF radio is circularly polarized, each antenna is phased to the next by 90 degrees in a circular order.

Figure 30 shows the final design done using the harness function of Solid Edge. The function has a built

in radius limit where the required minimum radius can be used as an input. In most cases coaxial cables

were solitary at their mounting points, meaning that the tie-wraps only held the coaxial cable and no other

wires at their tie-points. This strategy was adopted since the coaxial cable assembly came very late in the

satellite assembly, preventing other tie-wraps from being closed if it was to be shared with the coaxial

cable. The routing of the ISL radio, in the case of CanX-4/-5 resulted in the final assembly order of

assembling the X panels before the Y panels since access to the SMA connectors on both the ISL radio

and the panels needed to be performed once the X panels were in place. The ISL coaxial cable routing is

also shown in Figure 30.

62

Figure 30: Coaxial cable routing design for UHF and S-band (Left) and ISL (right)

4.4.6 Computer board connector placement

Having had the opportunity to design the harness while the computer boards were not fully designed,

some flexibility was available for the author to provide the best possible connector location for the power

board and payload boards. The OBCs had already been designed at the time of the author’s involvement;

hence its connector layout was fixed and used as a reference. For the Power Board (power board), the

choice of connector locations followed a particular order of reasoning. The design of the power board

connector layout was first designed with the position of the OBC connectors in mind, such that the power

connector on the power board for a specific harness was in proximity to the OBC data connector. This

ensured that all wires for each component harness would originate from a common area on the computer

board stack, ensuring a more organized design with less wires overlaps. Figure 31 demonstrates this idea

where the power connector for the reaction wheels, for example, is near its corresponding data connector

on the HKC. The other important consideration was the location of the component with respect to the

board stack. This positioning strategy is evident for the location of the panel connectors and was

followed for the BCDR connectors among others. Since the X panels were assembled before the Y panels

such that the ISL coaxial cables could be accessed from the Y side, it was required to place the X panel

connectors on the Y sides of the power board. The same approach was used for the Payload Computer

(PAYC) layout design, with the first consideration being the payload location. A complete revisit of the

payload harness needed to be performed following the replacement of the single CanX-4/-5 payload

computer named Formation Flying Computer (FFC) by a dual board design, which will be explained in

further detail in Section 4.5.3. Figure 32 shows the payload harness distribution used to determine the

optimal position for connectors on the Payload Power (PAY-POW) daughter board. A similar diagram

was produced by the author for AISSat-1 and BRITE.

63

Figure 31: Board stack showing connector layouts for BRITE

Figure 32: CanX-4/-5 payload harness over board stack with new dual board design

64

4.4.7 Integration procedures documentation

In order for the integration harness to be repeatable and achievable by a third party, it was necessary to

devise means of clearly presenting the harness mounting points and routing scheme. The routing of

harnesses is defined by the path it takes from tie-point to tie-point. It was then necessary to properly label

each tie-point such that it may be referred to in the routing instruction. Figure 33 shows the tie-point

representation method used in the GNB assembly procedures document. Many different sources of

information are present in this illustration. Each separate tie-wrap is identified such as T34. The color of

the tie-point label box represents the predominant harness it holds. The colored circles under the box

represent other harnesses it holds, according to the harness color legend provided. As seen in the lower

right legend, a tie-wrap orientation scheme is also presented such that the assembler may orient the tie-

wrap buckle properly. Every connector number is shown with the scheme for CanX-4/-5 shown in the

upper right corner. Finally, this illustration also shows heat shrink location for harness protection.

Although the model shown is BRITE, the tie-wrap color coded labels are generic for all missions.

Figure 33: Tie-point representation system

65

The same labeling strategy was used for the tie-wraps on the panels and for the AISSat-1 tie-wrap

designations. Throughout the assembly procedures, these tie-point numbers are used to describe the

routing sequence to be followed by the different harnesses. The specific instruction resembles the

following: “Have the two wires from pins 1 and 2 of the J4 (ADCC) connector go through tie-wrap T28,

T50, T49, T48, T44 and T42”. To summarize all harness routings, a full routing table is presented in the

appendix of the assembly procedures as represented in Figure 34. This table follows the same color code

as was used in the tie-wrap designations above. It is separated in three sections: one for generic harnesses

and two for mission specific harnesses. In the origin and destination columns, the pin-out, connector used

and number of wires is mentioned. As shown in the figure, splices are also presented with separate rows

for each splice. A similar table was developed by the author for the AISSat-1 harnesses.

Figure 34: Routing table for all harnesses (for representation only – text not visible)

For the more complex harnesses containing several connectors and involving several computer boards,

harness specific diagrams were presented such as the CanX-4/-5 payload harness which was shown in

Figure 32 and the sun sensor harness shown in Figure 35. In these schematics, the tie-wraps used as well

as the physical locations of each connector and tie-wrap are represented. The board stacks for each

mission as was shown in Figure 31 are also provided in the appendix of the assembly procedures

document for better visualization of connector locations. These diagrams, combined with the routing

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tables and the tie-point designation illustrations provide a complete set of tools to properly integrate the

wiring harness to the GNB satellites.

Figure 35: Sun Sensor harness schematic

4.5 Wiring harness critical design changes

This Section will present critical design changes which resulted in a revision of the harness design and its

integration. These changes occurred in chronological order, as can be seen in Table 8 in Section 3.3.4

describing the revision history of the GNB assembly procedures document [25].

4.5.1 Redundant use of OBC’s

Early on in the author’s involvement in the harness design, it was decided by the computer hardware

designers to add the capability to invert the roles of the OBCs such that a certain level of redundancy was

available. This meant that the House Keeping Computer (HKC) could also take on the role of the

Attitude Determination and Control Computer (ADCC) in the event that the ADCC was not behaving as

required, or vise-versa. The implication from this change was the necessity for the harness from each

component to have a connection to both OBCs. This was reflected by the use of dashed lines in Figure

25. A little more than 20 wires were added in order to implement this inter-OBC architecture, with

variations in this number depending on the mission. The general strategy for adapting this new

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architecture was to maintain the original wiring harness routing plan before the added wires and simply

find an available path for the additional wires.

4.5.2 Stage change of harness assembly

An important consideration during the writing of the assembly procedures, as described in Section 3.3.4,

was deciding at which time and to which tray the harness should be assembled first. The deciding factor

was determining which strategy was the most time efficient regarding project progress. Originally, it was

planned to assemble the wiring harness in the full assembly Section in order to simplify the process.

However, certain components such as the UHF and S-Band radios were initially not available during the

first flight-like test assembly. Because the radios cannot be assembled to the GNB once the payload is

assembled to the trays, it was illogical to proceed to full assembly as it would need to be redone once the

missing components were available. Thus, the harnesses were partially assembled to either tray with the

intent to complete the integration of the harness once the radios were made available and the full

assembly could be performed. As mentioned in Table 8 in Section 3.3.4, Revision 1.3 of the GNB

assembly procedures reflected this change in integration strategy. Figure 36 shows the partial harness

assembly to the +Z tray.

Figure 36: Partial harness assembly to the +Z tray

For the final version of the assembly procedures, Revision 1.5, it was decided to move the assembly of

the radio harness to the –Z tray, instead of the +Z tray. The main reason for this change in strategy was to

have the radio harness associated to the –Z tray such that the tie-wraps on that tray would not need to be

cut if later disassembly of the full satellite was required. This was necessary as these tie-wraps become

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difficult to replace once the S-Band radio is in place. It also permitted the assembly of the separation

switch at the sub-assembly level, rather then the assembly level.

4.5.3 Single to dual board design of CanX-4/-5

In order to make use of the already designed OBCs for the HKC and ADCC, it was decided to make use

of a third OBC as the payload computer for CanX-4/-5 rather than make use of a dedicated computer

which would have included both computing and power control for the CanX-4/-5 payload, called the

Formation Flying Computer (FFC). This new board would have required much designing effort, which

was not a viable programmatic option. For power regulation, a separate daughter board called Payload

Power (PAY-POW) is used in order to provide power to the dedicated OBC for CanX-4/-5 called the

Payload computer (PAYC). The PAY-POW is by comparison to the originally planned FFC, much

simpler to design. Figure 37 shows the original layout of the single board FFC and the new dual board

design PAYC and PAY-POW. The BRITE payload computer is shown for comparison.

In order to reduce required design effort, it was desired to make the harness as generic as possible

between missions. The connector layout of the original FFC designed by the author reflected this

strategy, where connectors for the power input and for the connection between the payload computer and

the OBCs were maintained at the same location. The connector for the payload harness was also located

at the same location on the payload board. Even though the components were different for both missions,

the harness routing was almost identical thanks to this convenient board layout. For example, the GPS

harness took the same path as the Star Tracker harness and the CNAPS harness took the same path as the

Telescope harness. The change to the dual board approach however required a revision of the harness

design and a departure from the mentioned generic strategy. Nevertheless, the choice of orientation of the

PAYC and PAY-POW, as well as the connector layout on the PAY-POW attempted to maintain the

harness routing as close as possible to the original routing with the FFC. Because of the additional link

required between the PAYC and the PAY-POW, a total of 25 wires were added. The routing for the new

harness needed to be carefully decided since the area available between the trays and panels for the wiring

harness was very limited. Figure 32 in Section 4.4.6 showed the final routing design based on the new

dual board approach.

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Figure 37: Payload computers connector layouts (Left: BRITE, Center and Right: CanX-4/-5)

As was shown in Figure 6, AISSat-1 also makes use of a third OBC for its PAYC as is the case for CanX-

4/-5. However, it does not require the PAY-POW daughter board. Although the harness from BRITE to

CanX-4/-5 was not completely generic due to the difference between a single to a dual board, making use

of the exact same computer board (the OBC design) between CanX-4/-5 and AISSat-1 provides the

ability to be even more generic in terms of the harness design. It is recommended, from an AIT point of

view, that future GNB missions make use of this same PAYC and the optional PAY-POW for the harness

to be as generic as possible.

4.6 Alternative Integration Concepts

From experience gained on the designing of the wiring harness on the GNB missions as well as

observations of other approaches used by other satellite missions, different alternative AIT concepts were

researched by the author. The most relevant for this thesis is the use of flexible circuits as an alternative

to conventional wiring for satellite integration.

4.6.1 Flexible circuits

This section will cover consideration of flexible circuits as an alternative to conventional wiring harness

for the next generation of GNB satellites. As will be described, careful use of flexible circuits could offer

numerous design improvements over conventional harnesses, especially concerning AIT aspects. As this

thesis covers AIT of the GNB, this aspect will also be the focus for flexible circuit considerations. This

coverage of flexible circuits is a conceptual application and therefore qualitative in nature and is

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presented merely to introduce its potential use in future missions. More detailed considerations and

careful discussions with flexible circuit manufacturers will need to be performed before deciding to take

this route. In particular, many different materials are available for conductors, insulators, adhesives and

finishes. Also, various options are available such as single sided layer, multi layers or hybrid rigid flex

circuits. Finally, different manufacturers have different capabilities, such as selective bonding and

preformed bends. All of these considerations will need to be evaluated for proper use of flexible circuit

technology.

4.6.1.1 Flexible circuits description

A flexible circuit is an array of conductors bonded to a thin dielectric film. A dielectric film consists of

non-conducting material such as Dupont’s Kapton material. The formal introduction to modern flexible

circuit technology is attributed to Frederick Seymour who published a generic patent to deal exclusively

with the manufacture of flexible circuits in 1927 [34]. However, only since the Second World War has

flexible circuits been used in the industry. According to a recent book on flexible technologies [35],

flexible circuits remain one of the fastest growing electrical interconnection products. Aerospace

missions such as NASA’s Mars Pathfinder mission used flexible circuits with Dupont’s Pyralux material

for the dielectric film10 to take advantage of the lightweight and reliability aspects of flexible circuits.

The handbook of flexible circuits [34] specifies two general application types for the selection of flexible

circuit material and manufacturing process selections. The first application is flex to install applications.

The industry generally calls this flex-to-fit. As the name describes, this application is where the flexible

circuit will remain static once installed. Its flexibility is used to help in its assembly. The second type is

dynamic flex applications. This type can be found in printers and scanners where repeated flexing

cycling will occur. Although the envisioned use on the future GNB would be of the first type, vibration

considerations during the launch of the satellite will produce cycling on the flexible circuit as if it was of

the second type. The most common manufacturing strategy for a dynamic flexing application is a treated

Rolled and Annealed (RA) copper for the conductor array which is sandwiched between polyimide films,

such as Kapton which offer superior cycling capability and a wide range of operating temperatures. RA

copper is formed by heating and mechanically rolling ingots of pure copper to the desired thickness.

10 http://www2.dupont.com/Kapton/en_US/uses_apps/aero/mars_rovers.html

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4.6.1.2 Flexible circuits application benefits study

The benefits and limitations of flexible circuits must be carefully considered before make use of it. Table

16 lists the relevant benefits and limitations over a conventional point-to-point wiring harness. As can be

seen, many benefits are offered with this technology, but careful design considerations are needed in

order to reduce the effects of the listed limitations. The limitation concerning limited current can be

easily overcome by multiplying the amount of conductor traces, which is easily accommodated by the

closely spaced traces. Since nanosatellites such as the GNB deals with relatively small currents, this is

less of an issue. The features listed are derived from [34] and personal assessments based on AIT

experience.

Table 16: Benefits and limitations of flexible circuits for GNB integration

Benefits Limitations

• Reduce workmanship related errors associated with the construction of harnesses

• Greatly simplify and accelerate assembly • Significantly reduce weight and volume • Improved reliability due to reduced handling,

number of connectors, reduced mass and stress on connections and joints

• Uniform electrical characteristics because of consistent spacing and orientation of conductors

• Flexibility in integration strategies • Allow up to a billion flexing cycles (high

resistance to vibration) • Ease of cleaning, storage and inspection • Consistent interconnect assembly

• Higher non-recurring cost in design stage • Low tolerance to changes and alterations • Limited current carrying capacity • RF signals susceptible to EMI

To overcome the limitations of flexible circuits concerning high frequency signal EMI issues, The NASA

stripline or microstrip technologies could be used in the design of the flexible circuits. These

technologies incorporate the insulating approach of coaxial cables by sandwiching the conducting layer

by two ground planes. With the current GNB design however, making use of traditional coaxial cables

for RF transmission would most likely be a simpler and cost effective approach. For this reason, flexible

circuits will only be considered for power and telemetry use in this study.

4.6.1.3 Potential design with flexible circuits

The key to making good use of flexible circuits is capitalizing on its offered benefits and reducing the

effect of its limitations. The proposed use of flexible circuits on the next generation of the GNB would be

to replace the wiring harness such that a more reliable, lightweight, consistent and simple interconnect

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strategy be used. The concept envisioned consists of a large branching sheet of flexible circuit which

would be folded around the structure to join every component with their respective connectors on the

computer boards. In terms of AIT requirements, this would improve assembly time and reliability and

permit an easier access to different components, including the central payload, without the need to

disassemble the interconnecting medium. The ease of cleaning of flexible circuits would be a major

advantage over wiring harness, as harnesses tend to be difficult to clean properly, as was discussed in

Section 3.3.2. With wiring harness, it is very difficult to estimate and model the EMI characteristics of

the harness as the spacing between conductors and their orientation cannot be consistently reproduced

during assembly. With flexible circuits the distance between conductors are well known and controlled

such that EMI can be modeled and predicted. With small conductor spacing and appropriate ground trace

locations, superior EMI reduction is expected as compared to conventional wiring harness [34].

As was experienced during the development of the GNB computer hardware and mentioned in Section

4.5, the computer boards underwent frequent changes, even at advanced stages in its development. The

limitation of flexible circuits in terms of being intolerant to changes is therefore a great concern when

considering these frequent computer board design changes which are expected in a satellite prototype

design. In particular, any relocation of connectors on the computer boards could potentially require a new

flexible circuit to be developed. The approach suggested in order to avoid this limitation would be to

design the computer boards with flexible circuits in mind, such that a single large connector would be

used. This connector would be conveniently located for easy routing and include extra connection lines

to permit changes on the computer board without affecting the flexible circuit. The flexible circuit design

would make use of traces with identical electrical properties so that their function can be interchanged.

The interconnect system could be made modular such that a main flex circuit would leave from the

computer boards, bridge across to the other tray and have separate component specific flex circuits

connect to it. As an alternative, a “daisy chain” systems architecture could be implemented such that each

trace would be dedicated to a specific signal type (such as signal transmit and receive) which would leave

the computer boards and then branch to the separate components. Communication to these components

would be ensured through their dedicated address. This approach would likely be less susceptible to

changes since the different signal types required will be established early. For that reason, the entire

interconnect system could be build from a single flexible circuit. This would also greatly reduce the

amount of conducting lines to the computer boards, thereby increasing reliability. The same approach can

be used for the connection of the redundant OBCs, currently achieved using splices. Splices, as opposed

to the trace branching used on flexible circuits, adds undesired electrical resistance and power loss.

Finally, the flexible interconnect circuit could be designed such that all bridging between trays be

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performed at a single location, currently impossible due to space constraints (see Section 4.4.2). This

would allow easy access to the payload, where both trays could be disassembled from the payload

structure without the need to disassemble the flexible circuit. Both the modular and daisy chain concepts

are shown in Figure 38.

Figure 38: Modular (left) and daisy-chain (right) concepts for use with flexible circuits

Manufacturing features, such as those shown in Figure 39, can be exploited to improve the flexible circuit

design. Holes cut in the flexible circuit could allow simple assembly to the structure with the use of

screws. Stiffeners could be added at mounting points to improve rigidity and protection against vibration.

Stiffeners could also be added when the flexible circuit must bend around a sharp corner in order to avoid

flexing at these bends. Another option could be to add pre-formed bends as shown or incorporate large

corner radiuses directly in the structure. A wide range of connectors specifically designed for flexible

circuits are available. Innovative use and placement of these connectors could reduce flexible circuit

bend radius and permit easier assembly. Certain components, such as sun sensors for example, could be

integrated directly into the flexible circuit by the use of rigid-flex technologies, which involves bonding

the internal layers of the flexible circuit to an external board layer for component surface mounting.

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Figure 39: Flexible circuits manufacturing capability examples11

4.6.1.4 Suggested implementation strategy

In order to minimize the design cost and permit some tolerance to change during the overall development

progress of the satellite, the following implementation strategy is suggested, according to the various

assembly phases presented in Section 3.3.1. First, during the design phase of the flexible circuit, CAD

models should be used for visualization. The sheet metal function in the Solid Edge CAD software could

be of particular use to visualize both the bent and flattened view of the flex circuit. During the fit check

assembly phase, the physical structure should be used with a simulated paper printout of the flexible

circuit to validate the design. During the first test assembly, a mock flexible circuit consisting of only the

dielectric film with bonded connectors and stiffeners could be used to verify that its assembly with all

satellite components can be performed in the planned order. For the flight-like test assembly, a first full

printed flexible circuit should be produced for final validation and system functional testing. This test

version of the flexible circuit would be made of a single piece with a provision for modularity by

providing intermediate mounting points for connectors in the event that a portion of the flex circuit would

need to be corrected. The corrected portion would be composed of a secondary flexible circuit piece

connected to the original flexible circuit. As an alternative, the secondary flex circuit piece could be

directly soldered to the original piece at a planned correction mating point. When flight assembly is to be

performed, it is expected that the satellite design will be mature such that the flexible circuit will no

longer require changes. At this point the final flight version of the flexible circuit would be produced

from a single piece without the intermediate connector mounting or solder joining points. This strategy,

11 These manufacturing capabilities are offered by www.flexiblecircuit.com

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from test to flight flexible circuit, is shown in Figure 40. Following a progressive implementation

approach such as this will minimize the costs associated with the manufacturing of the flexible circuit and

allow corrections to be made at every level before the final product is produced.

Figure 40: Flexible circuit implementation strategy from first test version to flight version

4.6.1.5 Flexible circuit implementation recommendations

As suggested in the handbook for flexible circuits [34], it is recommended to make full use of

manufacturer design expertise for guidance in the design phase. Many subtle but important design

differences, such as the choice of adhesive and dielectric material, can have serious impacts on its

tolerance to vibrations, for example. Many layout software tools are available for PCB designs. It is

important to understand these limitations of the design software when designing flexible circuitry. For

example, right angle turns in conductor traces, which are often used in PCB board designs, should be

avoided in flexible circuits as radiuses reduce stress concentrations during flexing.

If after careful assessment it is decided to use flexible circuits in the next generation of satellites at SFL,

the decision should be taken early in the design phase such that the structural design, the telemetry and

command system, and the board layouts be tailored towards its use. As was experienced with the design

of the wiring harness routing, its design in parallel with the board layouts permitted early changes, which

avoided potential problems during the satellite assembly. Necessary resources and time should therefore

be allocated for the proper design of a flexible circuit interconnection system. Although the non-recurring

cost during the design phase is greater for flexible circuits, the considerable time that will be saved during

the integration and assembly of the satellite could potentially reduce the overall cost of the mission and

significantly simplify the process. The author highly recommends the use of flexible circuits as an

elegant and practical alternative for AIT.

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5.0 GNB THERMAL TESTING

The author was involved with some aspects of testing for GNB missions, in particular with BRITE and

AISSat-1. At the component level, the solar cell Thermal Shock (T-Shock) test was conducted for the

BRITE mission. At the systems level, the author was involved with the design of the Thermal Vacuum

(TVAC) support structure that shall be used as a jig to hold both AISSat-1 and BRITE in the TVAC

chamber at the David Florida Laboratory (DFL) in Ottawa. The author will act as on operator for the

upcoming TVAC testing for both AISSat-1 and BRITE at DFL.

5.1 BRITE Solar cell Thermal Shock

This Section will describe the Thermal Shock (T-Shock) testing conducted for the BRITE mission’s solar

cells. A T-shock test is performed to both simulate extreme temperature changes in the event of a

tumbling satellite and to verify workmanship. A full inspection under a microscope was conducted before

and after the T-shock exercise to verify if any changes have resulted from the T-shock procedure. This

inspection was conducted by Mihail Barbu with assistance from the author. As described in the SFL

general thermal shock test procedures developed for the CanX-2 mission [36], a minimum change rate of

25°C / min is required to achieve desirable thermal shock. Temperature extremes should never exceed

the rated survival temperature limits of the component being tested. The solar cells themselves can

operate from -180°C to 95°C, whereas the bonding RTV silicon can operate in the range of -115°C to

260°C. A total of 25 cycles was conducted on all panels as was prescribed in the thermal shock

procedures document. The author produced a document describing the specific procedures used for the

T-Shock testing of the BRITE solar cells, as well as the achieved results [37].

5.1.1 Solar cell T-Shock preparations

In order to achieve the desired temperature changes of 25°C / min, the thermal mass of the test setup

needed to be sufficiently low. For that reason, two separate T-shock exercises were conducted with two

separate stacks of panels. The first stack was composed of the +X,-X and +Z panels, while the second

stack was composed of the +Y, -Z and –Y panels. Two thermal chambers where used with the following

settings: Hot Chamber: +100°C, Cold Chamber: -70°C. The initial temperature limits for the stacks

before being moved to the other chamber was +60°C and -20°C. For stack 2 (+Y, -Z ,–Y) these limits

where increased to +65°C and -30°C, in order to reach the desired rate of change of 25°C / min as stack 2

had a slightly higher thermal mass. As the chambers were set to more extreme temperatures, careful

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monitoring of the panel temperatures was necessary. In order to properly stack the panels without the risk

of damaging the solar cells, spacers, standoffs, screws and nuts where used. Figure 41 shows stack 1

(+X,-X,+Z). A similar stacking approach was used for the second panel stack. As can be seen, the solar

cells are well protected within the stack such that any objects that would be inadvertently dropped on the

stack would have less chance of damaging the fragile solar cells. A temperature sensor was bonded to

both top and bottom surfaces of the stack to get temperature reading during the T-shock process. The

center panel did not have a temperature sensor as its temperature was always bounded by the top and

bottom panel temperatures.

Figure 41: T-Shock panel stack 1 (+X,-X,-Z) assembly

Once the stack is complete, it is placed inside a static bag with two desiccant packs to absorb moisture.

Also, the use of a mini-vacuum removes the air from the bag, which is then sealed, which also reduces the

amount of humid air around the panels which could cause condensation on the solar cells.

5.1.2 Solar cell T-Shock procedures

The full detailed procedures for the BRITE solar cell T-shock can be found in [37]. Table 17 shows a

summary of the procedures used. The data collection software used permits logging of temperature of

both sensors and temperature rate changes. The actual T-shock cycling process occurs from steps 4 to 7.

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Table 17: BRITE solar cell T-shock procedures summary

T-Shock procedures

1. Connect the temperature sensor serial cable to the computer to log data 2. Set the hot chamber to 100°C and cold chamber to -70°C 3. Record time and place stack inside hot chamber for initial warm up until it reaches 60°C 4. Record time and place stack inside cold chamber and monitor temperature rate of change 5. Once stack is at -20°C, record time and transfer to hot chamber 6. Monitor temperature rate of change as temperature increases until it reaches 60°C 7. Repeat 4 to 6 for a total of 25cycles 8. Let stack cool to room temperature and stop data collection

5.1.3 Solar cell T-Shock results

The time of chamber transfer occurrence, maximum and minimum temperatures reached and maximum

and minimum temperate change rates reached were recorded during the T-shock process. Some

interruptions occurred due to a faulty wiring harness for the temperature sensors. Due to these

interruptions, some of the temperature change rates where lost, but data from the other uninterrupted

cycles seems to indicate that the rates were fairly consistent under similar conditions. The rates of

temperature change were averaging at approximately -26°C / min when performing cold cycles (step 4)

and at approximately +28 °C / min when performing hot cycles (steps 5 and 6). A temperature profile

curve for stack 1 is presented in Figure 42 for all cycles, with the red and blue lines representing the top

and bottom sensor respectively. The few appearing inconsistencies in the profiles are due to the

interruptions in the data collection causing distortions in the horizontal axis. The completion of 25 cycles

and the achieved temperature rate changes from the data obtained indicates that the T-shock process was

successful.

Figure 42: Temperature profile of stack 1 during T-shock for 25 cycles

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Once the T-shock was complete, a full inspection of the solar cells confirmed that they could withstand

the thermal shocking process. A single anomaly was observed under microscope and recorded, where a

solar cell had a small chip on an edge, where previously only a micro-fracture was present. As the chip

left no cracks, crack propagation was not a concern. This is believe to be caused by the handling of the

panel rather than caused by the temperature shocks considering that it was an isolated and non recurring

anomaly.

5.1.4 Solar cell T-Shock test recommendation

Although the process was successful, the method used for the BRITE solar cell T-shock could be

improved. The damaged harness for the temperature sensors causing interruptions in data acquisition

resulted from frequent transfer from the cold to the hot chamber and pinching by the door. Also,

frequently opening the doors to the thermal chambers created much condensation in the cold chamber,

which affected its performance to reach desired cold temperatures. It is recommended that a different

approach be used where no transfer of chambers is required. This would most likely require a single

thermal chamber which could quickly vary its internal temperature such that the proper temperature rates

on the test object could be reached. This was in fact the approach taken for the later AISSat-1 solar cell

T-shock testing which was performed in an external facility capable of performing thermal shock without

the need to transfer the panel stack. This approach also reduces the risk of damaging the solar cells from

physical handling of the panels.

5.2 Thermal-Vacuum Testing

The TVAC testing involves placing the fully assembled spacecraft inside a thermal vacuum chamber

which has the capability of both simulating vacuum such as in space and exposing the spacecraft to the

full range of expected temperatures to be experienced in orbit. By placing the spacecraft in a black

enclosed vacuum chamber with the temperatures of the walls nearing 77K, the predominant source of heat

to the spacecraft is through radiation from the InfraRed (IR) lamps, as is the case in space with radiation

from the sun. Although the radiation spectrums are different, the calculated heat flux is comparable.

A custom support structure is used to both hold the satellite and the IR lamps which will simulate the heat

flux of the sun on different faces of the satellite. Testing performed on the spacecraft includes a Worst

Case Hot (WCH) situation, with three panels illuminated by the IR lamps, and a Worst Case Cold (WCC)

situation, where only one panel is illuminated with the minimum expected flux intensity. While in these

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states, Long Form Functional Testing (LFFT) will be performed to ensure that the satellite behaves as

expected in all thermal conditions. The support structure makes use of long thin ropes to suspend the

satellite in order to reduce conduction between the satellite and the structure, thereby thermally isolating

it. The author was involved with a modification of the CanX-2 TVAC support structure and the full

design of the BRITE and AISSat-1 TVAC support structure.

5.2.1 CanX-2 jig structure modification

Although CanX-2 is not a GNB based mission, the lessons learned from its TVAC support structure

design was useful for the design of the GNB TVAC structure, described in the next Section. The initial

Thermal Vacuum testing for CanX-2 was performed at the Instrument Calibration Facility (ICF) at the

University of Toronto physics department during the summer of 2007. After having difficulties achieving

the desired cold temperatures and proper vacuum pressures, it was decided to perform a second full

Thermal Vacuum testing at the York University TVAC facility. Due to the different form factor of the

different chambers, the support structure, designed by Karan Sarda, needed to be modified by the author

in order to fit the new chamber. Simple modifications were required including shortening support legs,

narrowing the support plate and adding L-bracket fixtures. Figure 43 shows the modified CanX-2 TVAC

support structure in the York TVAC chamber with Cordell Grant, the CanX project manager.

Figure 43: Modified CanX-2 TVAC structure in YORK TVAC chamber

A desired improvement that was to be implemented on the new set up, based on experience at ICF was to

have the satellite suspension wires adjustable. Based on the assembly procedures for the York TVAC

testing [38], the satellite rests on an anti-sway platform during the tightening of these wires, which makes

it difficult to tighten all four wires evenly. The solution was to use simple commercial turnbuckles as

shown in Figure 44. These allowed the fine adjustment of the individual length of each wire so that the

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satellite is made level after the support plate is removed. It was also observed that the suspension

geometry permitted very little sway on the satellite.

Figure 44: Suspension strategy for CanX-2 using cables and turnbuckles

5.2.2 BRITE and AISSat-1 TVAC structure

As was done for the CanX-2 mission, the GNB satellites need to undergo full system TVAC testing. The

author was responsible for the design of the TVAC support structure for the BRITE and AISSat-1

missions. With guidance from the GNB thermal engineer, Karan Sarda, a set of requirements were

derived for this support structure. From the requirements, conceptual designs were produced for both the

York TVAC chamber and the DFL Thermal Vacuum chamber 4 (TV4). Initially it was not yet decided

which thermal chambers to use. From further discussions with DFL, the TV4 chamber was selected and

the conceptual and final designs were created, such that the frame may be fabricated and assembled at

DFL. This GNB TVAC structure supports the GNB spacecraft using a similar suspension approach as

was used for CanX-2. It also supports an array of IR lamps, an S-Band patch antenna and the wiring

harness bundles within the TVAC chamber. The final design document for the TVAC structure is

covered in [39].

5.2.2.1 Requirements

The author and Karan Sarda developed a set of requirements for the TVAC structure based on experience

from the CanX-2 frame. These requirements include considerations for the lamp placements, spacecraft

support strategy, structural support and for wiring harness. The requirements were created in order to

guide the design of the support structure such that the TVAC testing simulates the space environment as

close as possible while ensuring the safety of the satellite. Table 18 summarizes these requirements by

requirement types. The detailed requirements can be found in the TVAC support structure requirements

document [40].

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Table 18: GNB TVAC support structure requirements

Safety requirements

• There shall be minimal risk of collision or impact between the spacecraft and the assemblers during

assembly, which shall be verified through a mock assembly with the satellite model.

• There shall be minimal movement of the spacecraft within the chamber during the test such that the

satellite cannot come in contact with anything else than the support wires.

Frame requirements

• The support frame shall be able to support a GNB spacecraft, the required IR lamps, an S-band patch

antenna facing either Z panels and the wiring harness.

• The frame shall be designed such that there is minimum view factor blockage between the spacecraft

and the chamber walls.

• The frame should be assembled outside of the chamber.

• The frame shall be fastened to the chamber.

• The frame shall be grounded to the vacuum chamber.

• The frame should have high emissivity surfaces.

• The frame should be able to be disassembled easily for storage.

Spacecraft support requirements

• The spacecraft shall be centered about the axis of the cylindrical TVAC chamber

• The satellite shall be rigidly held in place while the suspension system is assembled

Lamp support frame requirements

• The lamps shall be arranged such that WCH (3 panels) and WCC (1 panel) can be simulated

• The lamps should be able to illuminate all faces

• The lamps shall be positioned such that they are aligned with the center of volume of the GNB cube

• The lamp to panel distance shall be less than the distance between the lamp and the VHF antenna

(AISSat-1) or the magnetometer boom

• The lamp filament shall be at least 10cm away from the spacecraft panels

• The lamp distance from the satellite shall be adjustable

• The lamp distance from the satellite shall be measurable to within 5mm

Wiring harness requirements

• The wiring harness shall be routed on the frame such that view factor blockage between the spacecraft

and the TVAC chamber walls is minimized

• The harness should be modular such that it can be assembled separate from the chamber and mated to

a separate harness in the chamber

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5.2.2.2 Conceptual designs

From discussions between Karan Sarda and the author, many different potential concepts were developed

with an effort to satisfy the requirements. Figure 45 shows some of these concepts, making use of an

AISSat-1 model for dimensional validity, since it contains the additional VHF antenna appendage which

must be taken into account.

Figure 45: Conceptual designs for the GNB TVAC support structure

Many different variants of a square type structure where developed while varying the satellite position,

lamp arrangements and suspension mounting points. Any such concepts where lamps were directly over

the satellite or too close to any of the appendages where then eliminated. Some concepts, as the image on

the left in Figure 45 shows attempts to minimize the number of lamp positions while still maintaining the

ability to simulated both Worst-Case-Hot (with 3 panels illuminated) and Worst-Case-Cold (1 panel

illuminated) situations. Because of the relatively small size of the available DFL chamber (3feet diameter

by 8feet deep), it was difficult to maintain an adequate lamp distance from the satellite with the cubic

frame concepts. The octagonal concept shown on the right, inspired by the support structure used for the

MOST mission, made better use of the space available at the expense of a seemingly more complex

assembly. However, from discussion with DFL and an assessment of the materials and fixtures they had

available for the frame construction, it appeared that an octagonal shape was the easiest shape for them to

construct.

5.2.2.3 Final design

Based on the octagonal concept described in the previous sub-Section, a final design was created by the

author such that it made use of the specific material available at DFL for its construction. Figure 46

shows the final design of the TVAC structure for both the AISSat-1 mission and the BRITE mission,

internally referred to as the “Thervoctatron”. More detail on its design and constituents can be found in

[39].

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Figure 46: Final design of the GNB TVAC support structure

As can be seen, the support structure is nearly identical for both AISSat-1 and BRITE. Since BRITE does

not contain the long VHF antenna, an extra lamp pair was added on the –Y face in order to provide more

simulation flexibility. The order of assembly follows the steps listed in Table 19. It makes use of DFL

designed extendable rails which permits the assembly of the structure outside of the TVAC chamber.

Table 19: GNB TVAC support structure assembly order

“ Thervoctatron” assembly procedures

1. Assemble the octagonal structure with 135° mounting brackets. 2. Assemble the octagonal structure to the four base tubes with 90° mounting brackets. 3. Rest assembly on the deployed chamber support rails and fasten with tie-wraps. 4. Assemble the two top turnbuckle supporting tubes. 5. Assemble the –Y lamp support tube with 90° mounting brackets. 6. Assemble all X, Z and +Y IR lamps and S-Band radio and route harnesses. 7. Assemble the turnbuckles to the top tubes. 8. Assemble the anti-sway platform sub-assembly (platform, tubes and 90° mounting brackets) to the

octagonal structure. 9. Rest satellite onto the anti-sway platform. 10. Fasten the suspension wires to the satellite and adjust tension with turnbuckles. 11. Route the satellite harness along the suspension wires. 12. Slide assembly with extendable chamber support rails into the TVAC chamber. 13. Remove the anti-sway platform. 14. Perform fine adjustments of the satellite (using turnbuckles) and lamp placements. 15. Assemble the –Y lamp tube assembly with lamps (BRITE only) and route harness. 16. Mate the wiring harness from the support structure to the chamber harness.

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In terms of the requirements, the design offers maximum safety to the satellite by avoiding lamps directly

over the satellite and allowing easy access to all other components during its assembly. By being highly

modular, the frame can be easily assembled outside of the TVAC chamber and later disassembled as well

as permitting easy adjustments to the structure. It offers very little view factor blockage along the Y axis

of the satellite and makes use of materials which have been used for other spacecraft missions at DFL,

thereby ensuring its heritage. As shown in Figure 47, the satellite is well centered within the chamber.

The frame can be easily fastened to the machined rails which make use of the mounting screws provided.

Figure 47: AISSat-1 TVAC structure shown in the DFL TV4 chamber12

12 DFL provided the CAD model for the TV4 chamber and tube mounting brackets

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6.0 AIT RECOMMENDATIONS AND CONCLUSION

6.1 General AIT Recommendations

Throughout this document, several recommendations for good practice and good design for AIT have

been made. This Section is intended to summarize these different recommendations and refer to the

Sections which contain the detailed description.

6.1.1 Assembly recommendations

In the assembly Section, several recommendations pertaining to good assembly practice and strategies

were presented. The overall objective of these recommendations is to ensure safe, repeatable and

effective means of conducting satellite assembly. The following point form list summarizes these good

methods, practices and design considerations.

A1. Make use of several test assemblies, such as fit checks, in order to detect issues early and be

more time efficient during flight assembly. Description of different assembly phases are

presented in Section 3.3.1.

A2. Proper documentation and revision control ensures repeatable assembly. See Section 3.3.2 to

3.3.5 for assembly documentation process used for the BRITE mission.

A3. Proper preparation for assembly sessions involves having a clear understanding of the task,

having a clean, organized working area, and ensuring proper storage of components. These

good practices are covered in Section 3.3.6.

A4. Constant communication between assemblers allows both peer review of execution

techniques and understanding each other’s intentions.

A5. A good surrounding awareness of people, components and tools is essential for the safety of

the satellite. This also involves making a conscientious effort to control and communicate

body movements. More detail can be found in Section 3.3.6.

A6. Safe assembly techniques include doing full inspection of components to be installed and

being meticulous in ensuring that the executed actions are correct. More examples are found

in Section 3.3.6.

A7. Never assume that a part is clean. Always do a full inspection and thoroughly clean

components before making use of them in the assembly. See Section 3.3.2 for the cleaning

process.

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A8. For any documented procedures, make use of past experience and own observations for

process documentation improvement.

A9. When designing GSE components, manufacturing considerations should be done such that

the parts can be machined with available equipment. Design strategies and machining

techniques can be found in Section 3.6.5.

6.1.2 Integration recommendations

In the integration Section, the author’s reasoning behind the design related to the interconnection medium

was presented. Also, an alternative AIT concept related to flexible circuits was introduced for future

missions. Recommendations derived from the integration experiences described in Section 4.0 will be

summarized in the following list.

I1. The definition of the integration strategy, including harness mounting points and routing

design should be conducted early so that modifications to the structure and computer board

layouts may be performed for optimal assembly and integration.

I2. The AIT engineer should always be consulted for any new computer board designs.

I3. Impacts of the integration strategy on the assembly order should be well understood.

I4. Disassembly of components, such as the central payload in the case of the GNB, should be

considered when designing the interconnecting medium.

I5. For ease of assembly, effort should be made such that the least amount of harnesses share

common mounting points. This is described in the mounting point strategy Section 4.4.3.

I6. Proper harness distribution and size estimation, along with physical test assemblies will

ensure wiring harness assembly feasibility. This process is explained in Sections 4.4.4.

I7. Proper documentation in the form of routing tables, diagrams and detailed instructions will

ensure integration repeatability. See Section 4.4.7 for integration strategy documentation.

I8. For future GNB based missions making use of conventional wiring harness, effort should be

made to maintain the same harness routing by making use of the PAYC and optional PAY-

POW computer boards as was done for CanX-4/-5 and AISSat-1. See Section 4.5.3.

I9. Flexible circuits should be considered for the future missions as an alternative to conventional

wiring harness. See Section 4.6.1 for flexible circuit description and potential use.

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6.1.3 Thermal testing recommendations

Some general recommendations can be derived from the work done on thermal testing and setup for the

T-Shock testing of the solar cells and the TVAC support structure for the fully assembled and integrated

satellite.

T1. A single thermal chamber capable of high temperature rates should be used for T-Shocking

such that no transfer of the test article from a hot to a cold chamber is necessary. See Section

5.1.4 for more detail.

T2. A good understanding of requirements and constraints should be established between TVAC

support structure designers and TVAC facility experts. See Section 5.2.2.1.

T3. Making use of available components as opposed to custom machined parts for the

construction of the TVAC frame can significantly simplify its assembly.

T4. Full assembly order should be considered when designing the TVAC support structure. The

assembly order and design of the GNB TVAC structure is described in Section 5.2.2.3.

6.2 Conclusions

This thesis has presented an overview of the aspects related to the assembly and integration of the GNB

based satellites. It has also presented work done on different thermal testing activities on the unit and

system level. As was evident in the frequent references between Section 3 and 4, assembly and

integration are very closely related and interdependent. The assembly sequence of the satellite is highly

dependent on the design of the wiring harness. The reverse is also true where the interconnecting medium

is subject to assembly requirements and constraints. A close relationship between the AIT designer and

the computer board engineers is essential for the proper integration and assembly of the spacecraft. This

interdependence is also true for every aspect of the satellite, supporting the need for a good understanding

of the different subsystems composing the spacecraft and good communication between the involved

engineers. A clear understanding of the subsystem requirements affecting AIT is necessary for the overall

success of the mission.

As was summarized in Section 6.1, this thesis made frequent recommendations concerning good

measures and practice, good design considerations and different potential approaches to solving AIT

challenges. The use of flexible circuits as an alternative is an example of a potential solution to meet the

integration needs of the missions of tomorrow. It is hoped that these recommendations will be followed

and developed further by the SFL engineers involved in AIT of future spacecraft.

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The experience gained by the author with his involvement in the AIT of the GNB missions lead to a much

better understand in the intricate interdependence between the different systems of a satellite. It permitted

exposure to different stages of spacecraft development from design to integration and testing. Finally, the

lessons learned from SFL allowed personal growth in terms of technical skills and improved satellite

design knowledge.

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