Preliminary System Development and Detailed Structural Design and Analysis for the Canx-7 Nanosatellite

Preliminary System Development and Detailed Structural Design and Analysis for the CanX-7 Nanosatellite

Fiona Singarayar

A thesis submitted in conformity with the requirements for the degree of Master of Applied Science Graduate Department of Institute for Aerospace Studies University of Toronto

Preliminary System Development and Detailed Structural Design and Analysis for the

CanX-7 Nanosatellite

Fiona Singarayar

Master of Applied Science

Graduate Department of Institute for Aerospace Studies

University of Toronto

2012

Satellites placed in LEO can remain there for an indeﬁnite period of time. To reduce the density of this orbit so as to avoid potential collisions with other satellites, the IADC has published a report that suggests any satellite in LEO should de-orbit within 25 years. CanX-

7 is a de-orbiting technology demonstration mission intended to help solve the global space debris problem. The work summarized in this thesis describes the author’s contribution to the

CanX-7 preliminary system development, as well as to the deployment detection and structural subsystems. Discussed herein are the challenges of carrying forward multiple designs in parallel and the factors and design trades that aid the decision-making process. This thesis not only presents the description of the ﬁnal design of the nanosatellite, but also the evolution of the spacecraft from when it was initially envisioned in 2010 to its current state at the time of this writing.

ii Dedication

For Amma - your unconditional love, tremendous support and unceasing prayers have shaped me into who I am today. Thank you for all that you do and for teaching me, “. . . More things are wrought with prayer than this world dreams of . . . ” I love you more than I could ever show.

iii Acknowledgements

I would like to thank God without Whom I would not be where I am today.

I would like to thank Dr. Robert E. Zee for giving me this truly incredible opportunity to be a part of a real space mission and to operate a real satellite! This has been an amazing experience and I’ve learned more about engineering in the past two years than in all of my undergraduate studies.

I would also like to thank the following people: Freddy—for answering all my questions, especially in my ﬁrst year when I didn’t know what the satellite was going to look like; Stephen— for his patience when answering all my mechanical-related questions; Cordell—for his feedback at PDR, and for thoroughly reviewing the CanX-7 solid model and providing suggestions to improve the structural design; Grant—for mentoring me and providing needed project management; Daniel—for teaching me about spacecraft operations, which has been a truly exciting part of my experience at SFL—and for even answering my phone calls when he was in India!

Paul, thank you for giving me a ride to school almost every morning. It made my commute tolerable and even a bit enjoyable with our bagel and coﬀee breakfast. It was always something that I looked forward to. Denis, thanks for your sense of humour, view on life and your readiness for a Tim’s run—rain or shine. I’m glad we got to extend our academic time together after undergrad for another two years. My time here wouldn’t have been quite as enjoyable without you. Thank you to my dear friend, Ash—for being like a sister to me. To my love and inspiration—Lawrence. Thanks for being a constant source of support, encouragement and love. You truly are the wind beneath my wings. Thanks for always being there for me and believing in me and helping me edit this thesis. I’m grateful to my sister, Nikita, for being awesome and understanding me like no one else can. Finally, thank you Amma for teaching me the important things in life—and also for making my lunch everyday as well as the freshly cooked meals that were ready when I got home late each evening.

iv Contents

1 Introduction 1

1.1 The Space Debris Problem ...... 1

1.2 CanX-7 Mission Objectives and Challenges ...... 2

1.3 CanX-7 Team ...... 3

1.4 Thesis Objectives ...... 4

1.5 Thesis Outline ...... 4

2 The CanX-7 Satellite 6

2.1 Drag Sail Payload ...... 7

2.2 Automatic Dependent Surveillance-Broadcast Receiver ...... 10

2.3 Structural Subsystems ...... 10

2.4 Attitude and Orbit Control Subsystem ...... 10

2.5 Power Subsystem ...... 11

2.6 Thermal Control Subsystem ...... 12

2.7 On-board Computer Subsystem ...... 12

2.8 Communication Subsystem ...... 12

2.9 Deployment Detection Subsystem ...... 13

3 CanX-7 Deployment Detection Subsystem 14

3.1 Deployment Detection Subsystem Requirements ...... 14

3.2 Imager Trades ...... 15

3.3 Integration with 3U Bus ...... 16

3.4 Field of View Analysis ...... 17

v 3.5 CanX-2 Operations ...... 18

3.5.1 Commissioning and Testing of Colour Imager ...... 18

3.5.2 Colour Imager Experiments ...... 19

3.5.3 Imaging Campaign Results ...... 19

3.5.4 Conclusion ...... 20

3.6 Current Status ...... 21

4 CanX-7 Preliminary System Development 22

4.1 Driving System Requirements ...... 22

4.2 SFL Spacecraft Platforms ...... 24

4.3 Initial Design Options ...... 26

4.4 Technical Feasibility Evaluation ...... 26

4.4.1 Ease of Manufacturability ...... 27

4.4.2 Volume Margin ...... 28

4.4.3 Mass Margin ...... 31

4.4.4 Energy Margin ...... 31

4.4.5 Flight Heritage ...... 34

4.4.6 Technical Feasibility Results ...... 34

4.5 Cost Evaluation ...... 34

4.5.1 Development Cost ...... 35

4.5.2 Beneﬁt and Value ...... 39

4.6 Preliminary Design Decision ...... 40

5 CanX-7 Structural Subsystem 41

5.1 Structural Requirements ...... 42

5.2 CanX-7 Preliminary Structural Design ...... 42

5.2.1 Structural Concept ...... 43

5.2.2 Structural Modiﬁcation ...... 45

5.2.3 Spacecraft Layout ...... 53

5.3 Design Evolution ...... 56

5.3.1 Drag Sails ...... 56

vi 5.3.2 Secondary Payload ...... 57

5.3.3 Power System ...... 61

5.3.4 Magnetometer ...... 62

5.3.5 Magnetorquers ...... 62

5.3.6 Imager ...... 63

5.3.7 Primary Structural Components ...... 63

5.4 Satellite Layout ...... 65

5.4.1 Layout of Main Tray ...... 65

5.4.2 Layout of −Z Panel ...... 66

5.4.3 Layout of Other Internal Components ...... 68

5.4.4 Layout of External Components ...... 68

5.5 CanX-7 Mock-up ...... 68

5.5.1 Prototype Structural Components ...... 68

5.5.2 Wiring Harness Design ...... 71

5.5.3 Current Status ...... 72

5.6 Assembly and Disassembly ...... 73

5.7 Mass Budget ...... 75

5.8 Structural Analysis ...... 76

5.8.1 Background ...... 76

5.8.2 Finite Element Modeling ...... 78

5.8.3 Boundary Conditions ...... 80

5.8.4 Modal Analysis ...... 81

5.8.5 Quasi-Static Analysis ...... 81

5.8.6 Hand Calculations ...... 82

5.9 Future Work ...... 85

6 Conclusion 86

Bibliography 88

vii List of Tables

1.1 CanX-7 Team ...... 3

3.1 COTS camera trade study summary, with signiﬁcant trades highlighted . . . . . 16

4.1 Preliminary mission and system requirements ...... 23

4.2 Technical feasibility evaluation matrix ...... 27

4.3 Preliminary mass budget for three candidate CanX-7 designs ...... 32

4.4 Energy budget for Design 1: GNB ...... 32

4.5 Energy budget for Design 2: Hybrid ...... 32

4.6 Energy budget for Design 3: Triple CubeSat (3U) ...... 32

4.7 Non-recurring engineering cost breakdown by labour time of activity ...... 36

4.8 Raw diﬀerential material cost break down ...... 37

4.9 Assembly, integration and testing diﬀerential cost break down ...... 37

4.10 Flat satellite (ﬂat-sat) cost break down ...... 38

4.11 Launch cost break down ...... 39

4.12 Total diﬀerential development cost breakdown ...... 39

4.13 Value of CanX-7 ...... 40

5.1 Mass budget comparison of conceptual and current designs ...... 76

5.2 Detailed mass budget of the current design ...... 77

5.3 Launch vehicle providers’ acceptance level test requirements ...... 81

viii List of Figures

2.1 CanX-7 solid model ...... 6

2.2 CanX-7 mechanical conﬁgurations ...... 8

2.3 Drag sail placement on two of SFL’s platforms ...... 9

2.4 Current design of the drag sail module ...... 9

3.1 Imager integration on 3U bus ...... 17

3.2 Preliminary ﬁeld of view analysis ...... 17

3.3 Portion of sail expected to be seen by the C329-SPI camera ...... 18

3.4 Image taken two minutes after penumbra by CanX-2 colour imager ...... 20

3.5 An image of Earth’s horizon taken by CanX-2 colour imager ...... 20

4.1 Three main SFL spacecraft platforms ...... 24

4.2 Drag sail module mounting scheme and star tracker layout on the GNB . . . . . 29

4.3 Design 2: Hybrid ...... 30

4.4 Design 3: 3U ...... 31

5.1 Structural concept of both CanX-2 and CanX-7 satellites ...... 43

5.2 Drag sail design and stacked conﬁguration ...... 44

5.3 CanX-2 and preliminary CanX-7 main tray ...... 46

5.4 Modiﬁed main tray and +X solar panel ...... 46

5.5 Modiﬁed −Z panel and arrangement with main tray ...... 47

5.6 Modiﬁed −Z panel and +X solar panel ...... 48

5.7 Modiﬁed X solar panel ...... 49

5.8 Modiﬁed −Y solar panel ...... 49

ix 5.9 GNB tray and radio cover ...... 50

5.10 Parts of the UHF radio ...... 51

5.11 CanX-7 magnetometer at PDR ...... 52

5.12 Printed torquer design at PDR on Z− panel ...... 53

5.13 S-band built into main tray ...... 54

5.14 Internal layout at PDR ...... 55

5.15 External view of the satellite (right) and 180◦ rotation (left) ...... 56

5.16 Evolution of the drag sail module integration scheme ...... 57

5.17 Post-PDR state of the secondary payloads and their mounting location ...... 58

5.18 Current state of the secondary payload and its mounting location ...... 61

5.19 Evolution of the power system ...... 62

5.20 Evolution of the magnetorquer design ...... 63

5.21 Main tray design at PDR (left) and current design (right) ...... 64

5.22 −Z panel design at PDR (left) and current design (right) ...... 65

5.23 Main tray layout at PDR (left), post-PDR (center), and current (right) . . . . . 66

5.24 Radio assembly at PDR (left) and current layout (right) ...... 67

5.25 Layout of current −Z panel ...... 67

5.26 Current external layout ...... 68

5.27 Solid model and prototype of structural parts for the CanX-7 mock up ...... 70

5.28 Parts printed on support material ...... 70

5.29 Mechanical ﬁt check and partial assembly of CanX-7 mock-up ...... 72

5.30 Wiring harness for CanX-7 mock-up ...... 72

5.31 Spacecraft-level assembly procedure for CanX-7 ...... 74

5.32 Finite element model and the imposed boundary conditions of CanX-7 ...... 80

5.33 First natural frequency of CanX-7 ...... 82

5.34 Satellite stress spectrum under the worst-case acceleration ...... 82

x Chapter 1

Introduction

1.1 The Space Debris Problem

Space debris is the collection of man-made objects in space that no longer serve any useful purpose. As the number of objects placed in Low Earth Orbit (LEO) increases, the risk of collision grows. Such collisions further increase the number of objects in high density orbits such as LEO. This domino eﬀect will lead to exponential growth of the orbital debris population if mitigating measures are not taken. Current missions are threatened by debris, which have the potential to end or disrupt operational satellites upon impact. This situation was witnessed in February 2009 when the non-operational Russian satellite Kosmos 2251 collided with the active U.S. satellite Iridium 33 [1]. The collision resulted in a large number of fragments that now pose a threat to other operational satellites.

Mitigating debris generation is vital for the continuation of safe operation on orbit. In order to address this concern of outer space pollution, the Inter-Agency Space Debris Coordination

Committee (IADC) has convened in 2007 and has published the Space Debris Mitigation Guide- lines report that suggests a maximum 25-year lifetime for a satellite in LEO after completion of mission or 30 years in total [3]. These guidelines serve to minimize debris generation and act as a safety measure for future space missions.

Due to these IADC policies, which are heavily emphasized in Canadian satellite missions by the Government of Canada, the capability of de-orbiting a spacecraft becomes a mission requirement. That is, failure of compliance by satellite developers—such as the Space Flight

1 Chapter 1. Introduction 2

Laboratory (SFL) at University of Toronto (U of T)—may result in denied licensing for trans- mission frequencies and/or remote sensing [4]. “De-orbit,” refers to the decrease in altitude of the satellite during re-entry into the Earth’s atmosphere, where the satellite will either burn up or return to the surface of the Earth within a short period of time [4]. The inability to demonstrate that a satellite will de-orbit within the 25 years post-mission may hinder projects developed by the microspace philosophy—an approach to designing small satellites on short schedules at relatively low cost. Without a dedicated de-orbit strategy, small satellites at higher LEO altitudes—160 km - 2,000 km above the earth’s surface—can remain in orbit for up to a hundred years [2]. Therefore, solving the de-orbiting problem is imperative for micro- and nano-class satellites as they may be rendered high-risk for future space missions without a successful de-orbit technology.

1.2 CanX-7 Mission Objectives and Challenges

CanX-7 (Canadian Advanced Nanospace eXperiment 7) is a nanosatellite currently in development at SFL—which builds cost-effective micro- and nano-class satellites—to demonstrate a de-orbiting technology using deployable drag sails. CanX-7 will be one of the first of its kind to demonstrate the de-orbiting of a nano- or microsatellite through the use of a de-orbit device: drag sails. NASA’s NanoSail-D was primarily a solar sailing demonstration mission showing the efficacy of de-orbiting a small satellite using solar sails. However, this technology is unlikely to be utilized as a drag device on operational nanosatellite missions since two-thirds of the

NanoSail-D satellite volume was used for solar sail stowage and the remaining one-third volume was used to support the solar sail [5]. Unlike the solar sail on NanoSail-D, the drag sails on

CanX-7 are primarily a de-orbit device that will be used on other missions whose main objectives are not to de-orbit themselves. These drag sails are intended to be added to future SFL satellites as the de-orbit device and without interfering with primary missions. The CanX-7 main mission objectives are:

• To demonstrate the capability of the de-orbiting technology—passive drag sails—to de-

orbit SFL nanosatellites as large as a 15 kg spacecraft with some scaling of the sail;

• To validate post-deployment attitude and de-orbit models to aid in the de-orbiting anal- Chapter 1. Introduction 3

yses of future SFL satellites; and

• To operate a secondary payload before demonstrating the de-orbiting technology.

There are many challenges associated with achieving these high-level mission goals. From a systems and a structural point of view, these challenges include:

• Uncertainty of the CanX-7 bus design during the early stages of the mission;

• Fitting the drag sails into the limited space of 100 × 100 × 340.5 mm—a triple CubeSat

(3U) form factor; and

• Fitting the avionics into the limited space of a triple CubeSat form factor.

1.3 CanX-7 Team

The CanX-7 spacecraft has evolved tremendously from its conceptual design stage in 2010.

Over the last couple of years, it has been worked on mostly by a team of graduate students—as it is mainly a student-oriented project—as well as SFL staﬀ engineers. Table 1.1 summarizes the staﬀ and students who have participated in the CanX-7 project to date.

Table 1.1: CanX-7 Team CanX-7 Name Position Period Responsibilities Project Management, Dr. Robert Zee SFL Director Sept 2010–Present Design Review CanX-7 Project Freddy Pranajaya Management, Payload Sept 2010–Sept 2011 Manager CanX-7 Project Management, Mission Grant Bonin Sept 2011–Present Manager Analysis Payload, Mission Barbara Shmuel Masters Student Sept 2010–Present Analysis Fiona Singarayar Masters Student Systems, Structure Sept 2010–Present Vincent Tarantini Masters Student AOCS, Payload June 2011–Present Jesse Hiemstra Masters Student Payload Sept 2011–Present Bryan Johnston-Lemke Electrical Engineer Power Nov 2011–Present Jakob Lifshits Computer Designer Software Jan 2012–Present Chapter 1. Introduction 4

1.4 Thesis Objectives

The objective of this Master’s thesis was to learn the principles and practices of designing and building micro- and nanosatellites under the microspace philosophy and to be involved in the full design cycle of the CanX-7 nanosatellite mission: from conceptual design to launch. At the beginning of the author’s work, CanX-7 was very early in its design and provided an ideal opportunity for involvement in the nanosatellite design process. The goals of this Master’s thesis were to:

• Gain an understanding of the microspace approach to designing nanosatellites;

• Develop an understanding of systems engineering;

• Perform systems engineering work for CanX-7;

• Learn the principles of designing a subsystem;

• Design the CanX-7 deployment detection subsystem; and

• Design the CanX-7 structural subsystem.

1.5 Thesis Outline

This thesis consists of six chapters describing the author’s contribution to the CanX-7 mission.

Chapter 1 has provided background information on the space debris problem and summarized what SFL is doing to address this signiﬁcant issue. Chapter 2 provides an overview of the current payloads and the subsystems that comprise the CanX-7 spacecraft. Chapter 3 discusses the process involved in designing the deployment detection subsystem for CanX-7. The experience of being a satellite operator is also discussed in this chapter in the context of how it guided certain decisions in designing the deployment detection subsystem. Chapter 4 describes the system development of CanX-7 in its early stages of design. This includes the preliminary budgets and cost evaluations of initial CanX-7 buses upon which design trades were based.

Chapter 5 details the development of the structural subsystem—from preliminary design to prototype to analysis. This treatment includes not only the ﬁnal design of CanX-7 but also the designs that were rejected or abandoned during the progression of CanX-7. Also included is the current state of CanX-7 as well as the work remaining to prepare the CanX-7 spacecraft for Chapter 1. Introduction 5 launch. Finally, Chapter 6 concludes the thesis with a summary of lessons learned and work done to date. Chapter 2

The CanX-7 Satellite

The CanX-7 satellite in its current form (Figure 2.1) is a triple CubeSat (3U) form factor—

100 × 100 × 340.5 mm—consisting of two payloads and seven subsystems, all of which are essential to mission success. The most up-to-date description of each of the payloads and subsystems of the satellite are outlined in the subsequent sections.

Figure 2.1: CanX-7 solid model

As CanX-7 has a 3U form factor, it will use SFL’s XPOD Triple as its separation system to be ejected into orbit. The XPOD Triple has no open faces. As a result, a few components on CanX-7 are deployables. During launch, the drag sails and the deployables are in a stowed conﬁguration (Figure 2.2(a)). Once in orbit, CanX-7 will operate a secondary payload provided by COM DEV Ltd. for approximately six months. During this phase, CanX-7 will be in a

6 Chapter 2. The CanX-7 Satellite 7 mechanical conﬁguration as seen in Figure 2.2(b). Upon the end of the Secondary Payload

Operation (SPO) phase, the drag sails will be deployed (Figure 2.2(c)) in order to de-orbit the satellite so as to demonstrate what is possible for future compliance with IADC guidelines. Once deployed, the drag sail is a completely passive device and does not need an active satellite for its function. This demonstration relies on atmospheric drag created by the drag sails to remove energy from the satellite. This would cause CanX-7 to decrease in altitude until eventually it is low enough to burn up in the atmosphere.

The novel technology, which will be ﬂown on a 3U form factor spacecraft, is intended to de- orbit various other classes of buses at SFL. Two of such buses include Generic Nanosatellite Bus

(GNB)—a 20 × 20 × 20 cm design—and Nanosatellite for Earth Monitoring and Observation

(NEMO) class satellites. This is possible because the drag sail design does not use a centralized deployment strategy. The modularity of the design allows versatility in its placement within a satellite bus. On a 3U form factor the drag sails will be stacked on top of each other, while on a GNB or NEMO class satellite the sails will be tiled. Both conﬁgurations are shown in

Figure 2.3.

The simultaneous use of four drag sails is proposed for use on CanX-7 to suﬃciently prove that the de-orbiting technology could de-orbit the reference spacecraft—a 15 kg nanosatellite as large as the NEMO bus—within 25 years of sail deployment from a maximum altitude of 800 km. The reference spacecraft is the heaviest SFL spacecraft that the drag sails are designed to de-orbit. Since CanX-7 has a 3U form factor with a weight of approximately 3.5 kg, it is expected to de-orbit in a much shorter period than 25 years and therefore accelerate the de- orbiting demonstration. By demonstrating the success of the de-orbiting solution on-orbit, the drag sail will earn ﬂight heritage, which will enable continued microspace activity at SFL. By using this de-orbiting technology on future SFL satellites, the laboratory is helping to mitigate the global problem of orbital debris.

2.1 Drag Sail Payload

To accomplish its primary mission objective—to demonstrate the de-orbiting technology applicable to micro- and nano-class satellites—CanX-7 will employ modular passive drag sails. Each Chapter 2. The CanX-7 Satellite 8

(a) Deployables and (b) Deployables deployed after ejection drag sails stowed in from XPOD and during SPO phase XPOD during launch

Figure 2.2: CanX-7 mechanical conﬁgurations drag sail is initially stowed within its individual module and deploys to an approximately 1 m2 area. The drag sail design has evolved signiﬁcantly over the course of this Master’s thesis. In its current form, the drag sail modules have a triangular wedge shaped footprint with a 3 cm height, as seen in Figure 2.4. The drag sail itself is folded and stored in a sail cartridge, which

ﬁts into the drag sail module. The sails are attached to the end of two tape springs that serve Chapter 2. The CanX-7 Satellite 9

(a) Drag sail modules stacked on a 3U (b) Drag sail modules tiled on a GNB

Figure 2.3: Drag sail placement on two of SFL’s platforms as booms to deploy the sail. The booms are wound around a reel inside the module and held in place by a module door, which in turn is held closed by a mechanism [14]. The sail deployment is based on the concept of stored mechanical energy that will be released from the tension in the tape springs when the module door is opened.

Figure 2.4: Current design of the drag sail module Chapter 2. The CanX-7 Satellite 10

2.2 Automatic Dependent Surveillance-Broadcast Receiver

The secondary payload that CanX-7 will carry is provided by COM DEV Ltd. In its most recent form, it is an Automatic Dependent Surveillance-Broadcast (ADS-B) receiver that will be operated for approximately six months. The main purpose of the payload is not to perform message detection, but to demonstrate the ADS-B technology on orbit. It will acquire ADS-B signals from aircraft to gain an understanding of signal characteristic on orbit as well as possible interference [4]. It will serve as a load bearing structure for CanX-7 as it is one monolithic piece weighing approximately 500 g with an outer dimensions of 96.5 × 98 × 83 mm.

2.3 Structural Subsystems

The design of the CanX-7 structural subsystem, with its 3U form factor, is largely based on

CanX-2 as described in detail in chapter 5. The 3U design consists of a one large tray and ﬁve panels (−Z, +X, −X, +Y , −Y ) which will be machined from magnesium, and one smaller magnesium tray for the UHF radio. The majority of components are mounted to the main tray and the −Z panel, which both have launch rails built into them. The launch rails and the cross braces on the +X and −X panels will carry most of the load experienced during launch.

Very few components are mounted internally to the +X, −X, +Y and −Y panels. These are

2 mm thick magnesium sheets and are mainly used as a structural shell to hold the rest of the satellite together. The exterior of these panels serves as mounting surfaces for the solar cells.

Details of this subsystem are provided in Chapter 5.

2.4 Attitude and Orbit Control Subsystem

The attitude and orbit control subsystem (AOCS) is responsible for pointing the satellite in favourable orientation during the Secondary Payload Operation phase and recovering it from any tumbles. Attitude determination on CanX-7 is achieved through a three axis magnetometer which has earned ﬂight heritage by being ﬂown on AISSat-1. However, unlike in this GNB satellite, the magnetometer on CanX-7 is mounted internal to the satellite. In addition to the magnetometer, the panel currents will also be used to aid with attitude determination. For Chapter 2. The CanX-7 Satellite 11 attitude control, CanX-7 will utilize a set of three orthogonal printed magnetorquers, custom designed at SFL, each of which produces a maximum dipole of 0.2 A·m2. The printed torquers will be used on AISSat-2, which is to be launched before CanX-7, and therefore will have earned

ﬂight heritage when CanX-7 is ready to launch. The torquers will be used to detumble the satellite after ejection and also in case of a spin up. They will also be used to orient the satellite in a favourable attitude prior to sail deployment to minimize any impact on communications.

2.5 Power Subsystem

The main purpose of the power subsystem is to provide enough electrical power to keep the satellite alive and functional during the entire mission. The CanX-7 power subsystem is composed of four components: solar coupons, battery, BCDR (Battery Charge/Discharge Regulator) and a power control system.

CanX-7 will have a total of eight solar coupons which are externally mounted to the panels.

They serve as the main source of power generation and will be used when CanX-7 is in the sunlit portion of its orbit. During eclipse, the satellite will use a rechargeable 4800 mA·h lithium-ion battery as the second source of power. To ensure the battery temperature is within recommended limits, a battery heater is included and the battery is attached to a battery thermal control board located on the BCDR. The BCDR, in addition to monitoring the temperature of the battery, sets the operating voltage of the solar arrays, monitors battery voltage and regulates the charge and discharge current to the batteries.

The power control system on CanX-7 is a relatively new design. The current power control system design is modular—it consists of two micro Switch Power Nodes (µSPNs) which provide power and ground to each of the loads; an InterFace Node (IFN) which serves the main functionality of the GNB power board; and a POwer System InterFace (POSIF) board which connects the rest of the satellite to the power system. The two µSPNs, the IFN and the POSIF together make up the power control system. Chapter 2. The CanX-7 Satellite 12

2.6 Thermal Control Subsystem

Although not currently fully mature, the thermal control subsystem is intended to be passive.

This will be accomplished through thermal tapes, material selection and component layout.

The main purpose of the thermal control subsystem is to keep satellite and its components within a temperature range acceptable for proper functionality. Heaters, as active measures for thermal control, are only used as a last resort for essential components unable to meet this temperature requirement through passive means. The I-DEAS software will be used to aid in the thermal modeling and analysis of CanX-7.

2.7 On-board Computer Subsystem

The CanX-7 satellite will only have one On-Board Computer (OBC)—the House Keeping Com- puter (HKC) used on other GNB spacecraft. This is responsible for communicating with the rest of the subsystems and relaying information between them.

2.8 Communication Subsystem

The CanX-7 communication system is full duplex and is comprised primarily of two radios—

Ultra-High Frequency (UHF) Receiver (Rx) and S-band Transmitter (Tx)—both designed and built in-house. Both radios have ﬂight heritage as they were ﬂown on AISSat-1, CanX-2 and

NTS.

The UHF uplink is the only way of sending commands and uploading scripts or software to the satellite. It operates in the amateur band and will use an antenna system consisting of four canted monopoles that will deploy after ejection from the XPOD. This arrangement and deployable mechanism, providing near omni-directional coverage, are well established as they have been used on CanX-2. On orbit, the UHF Receiver will always be in an ON state. This way the satellite is always “listening” and CanX-7 may be reset or power cycled from ground when a ﬁrecode is issued. The S-band transmitter will provide the downlink capability for CanX-7 by utilizing two patch antennas. They are mounted on opposite faces of the satellite to provide near omni-directional coverage over the full sphere of the satellite. The S-band transmitter Chapter 2. The CanX-7 Satellite 13 nominally operates between 32 and 256 kps.

After launch, CanX-7 will be commissioned and operated from the ground station located at

SFL. While the UHF Receiver and S-band transmitter are used to communicate with CanX-7 on orbit, the test port is used as a means to communicate with it on ground. The test port is a

115 200 bps serial UART that allows communication between the Ground Support Equipment

(GSE) and the HKC.

2.9 Deployment Detection Subsystem

Although not currently mature, the Deployment Detection Subsystem (DDS) is to include a Commercial Oﬀ-The-Shelf (COTS) camera to image the deployed drag sails. The design progression of the DDS is discussed separately Chapter 3 as this was one of the two subsystems worked upon by the author as part of the Master’s thesis focus. Chapter 3

CanX-7 Deployment Detection Subsystem

Given that CanX-7 is intended to be the ﬁrst demonstration of SFL’s de-orbit sail technology, it was considered highly desirable that the spacecraft include a Deployment Detection Subsystem

(DDS). This section summarizes the deployment detection subsystem requirements and includes a trade study on the diﬀerent candidate cameras that were considered as the imager for drag sails of CanX-7. The author’s experience as a satellite operator is also discussed in how it relates to the DDS. Finally, a preliminary solution is presented and the decisions that led to its current state are also discussed.

3.1 Deployment Detection Subsystem Requirements

The following are the applicable system requirements for the deployment detection subsystem [11]:

SYS-22 The system shall gather engineering telemetry that can be used to detect or infer sail

deployment; and

SYS-23 The system should be capable of determining the percentage of sail deployment to

within ±10%.

Derived from these high-level system requirements and using a bottom-up approach— requirements arising from what is feasible to design—is the document, “CanX-7 Deployment De-

14 Chapter 3. CanX-7 Deployment Detection Subsystem 15 tection System Requirements and Veriﬁcation Matrix” [10]. The relevant requirements from [10] that were used to assess the camera include the following:

DDS-1 The deployment detection system shall capture an image of the sail after it has been

deployed;

DDS-2 The deployment detection system shall have a mass of no greater than 90 g;

DDS-3 The deployment detection system shall ﬁt within 23.5 × 56 × 28 mm;

DDS-4 The deployment detection system shall have an average power consumption of no

greater than 500 mW ;

DDS-5 The deployment detection system should have a ﬁeld of view (FoV) of at least 90◦;

DDS-6 The deployment detection system shall have an exposure time that can be controlled

manually;

DDS-7 The deployment detection system should have an operational temperature range of

−20 to +60 ◦ C and a survival temperature range of −40 to +80 ◦ C; and

DDS-8 The deployment detection system shall use asynchronous serial, I2C or SPI to commu-

nicate with the House Keeping Computer (HKC).

3.2 Imager Trades

The deployment detection subsystem is to include a Commercial Oﬀ-The-Shelf (COTS) camera which will image the sail directly. In order to choose a particular COTS camera for the CanX-7 imager, the deployment detection subsystem requirements presented in Section 3.1 were used.

Cameras that were considered include:

• VGA camera cube (from OMNIVision);

• Fully steerable wireless micro-camera (from Ingegneria Marketing Tecnologia);

• C329-SPI Camera Board (from Electronics123);

• SRV-1 Blackﬁn Camera (from Surveyor Co.); and

• µCAM Serial JPEG camera module (from 4D systems).

A summary of how these COTS cameras compared to the deployment detection subsystem requirements is presented in Table 3.1. The ‘deal-breaker’ for each unit is highlighted and is further explained below. Chapter 3. CanX-7 Deployment Detection Subsystem 16

Table 3.1: COTS camera trade study summary, with signiﬁcant trades highlighted

COTS Camera Fully Steerable SRV-1 µCAM Serial VGA Camera C329-SPI Criteria Requirement Wireless Blackﬁn JPEG Camera Cube Camera Micro- Camera Module Camera Mass [g] DDS-2 – 500 – 36 10 Dimensions [mm] DDS-3 2.8×3.2×2.5 70×70×130 20 × 28 × 18 50 × 60 × 40 32 × 32 × 35 Power usage [mW] DDS-4 0.125 0.48 0.064 0.145 0.064 FoV [◦] DDS-5 diag; 64 – diag; 100 diag; 100 diag; 120 Exposure control DDS-6 automatic – automatic manual automatic Temp. range [◦ C] DDS-7 −30 to +70 −20 to +60 −20 to +70 −20 to +70 −10 to +70 Data interface DDS-8 MIPI – SPI UART TTL or RS232

The VGA Camera Cube would be ideal for the purpose of imaging the sail since it is extremely small, allowing for multiple camera use to image all four sails. However, the its data interface—MIPI (Mobile Industry Processor Interface)—is not compatible with the interface required on GNB and therefore cannot be further considered. The Fully Steerable Wireless

Micro-Camera is too heavy to be considered as the target mass for a 3U bus is 3.5 kg. The problem with the SRV-1 Blackﬁn Camera is volume—the camera will not ﬁt within a 3U platform. Likewise, the µCAM Serial JPEG Camera Module from 4D Systems is also larger than the allocated volume. The C329-SPI Camera is able to meet all but the exposure control requirement (DDS-6). This issue could be addressed by overriding the auto exposure function but further investigation would be required. As a result, during Preliminary Design Review

(PDR) it was intended that CanX-7 should use the C329-SPI Camera Board, with further investigation into modiﬁcation required for exposure control.

3.3 Integration with 3U Bus

The imager allocation for the CanX-7 during the preliminary design was on the main tray as shown in Figure 3.1(a). There is a lens cutout on the tray through which the imager can view the sail, and another cutout for imager wiring. These measures are necessary because the main imager cutout will not be accessible after the UHF casing is mounted on to the main tray. The lens cutout could be increased by 5.5 mm in case the imager needs to be tilted (Figure 3.1(b)).

With a C329-SPI Camera, this corresponds to a maximum tilt of 10◦. Chapter 3. CanX-7 Deployment Detection Subsystem 17

(a) Imager cutout on main tray (b) Close up view of imager location

Figure 3.1: Imager integration on 3U bus

3.4 Field of View Analysis

The C329-SPI Camera has a ﬁeld of view (FoV) of 63◦ × 78◦, which corresponds to a 100◦ FoV along the diagonal. If the camera is rotated along its lens axis, the diagonal FoV could be used.

Conservatively assuming a 90◦ FoV, the C329-SPI Camera could image 0.8 m2 of the sail when mounted perpendicular to the tray as seen in the left side of Figure 3.2(a) and 0.9 m2 when mounted at a 10◦ tilt (right side of Figure 3.2(a)).

With a worst case scenario of a 63◦ FoV, the camera could image 0.36 m2 of the sail if it is mounted perpendicular to the tray (left of Figure 3.2(b)) and 0.51 m2 when mounted at a 10◦ tilt (right side of Figure 3.2(b)). A top view of the sail in the latter conﬁguration is shown in

Figure 3.3(a). The expected picture to be captured on-orbit would look similar to Figure 3.3(b) with part of the sail seen at a 10◦ tilt against the backdrop of space.

(a) Imager with a 90◦ ﬁeld of view (b) Imager with a 63◦ ﬁeld of view

Figure 3.2: Preliminary ﬁeld of view analysis Chapter 3. CanX-7 Deployment Detection Subsystem 18

(a) Top view of the sail (b) Expected picture to be capture on-orbit

Figure 3.3: Portion of sail expected to be seen by the C329-SPI camera

3.5 CanX-2 Operations

During ﬁve weeks of imaging experiments performed using the colour imager of the CanX-2 main on-board computer (MOBC), the author gained insight into factors aﬀecting DDS design.

This section details the design, testing, execution and analysis of these imaging experiments and their relevance to the deployment detection subsystem.

3.5.1 Commissioning and Testing of Colour Imager

In June 2011, after the very first commissioning of the CanX-2 colour imager, several test images were taken on orbit. However, they were either over- or under-saturated. To better understand the influence of the imager parameter settings on these results, testing was carried out on the CanX-2 flatsat. All parameters ultimately influenced the exposure time. The imager functions such as “Time Delay,” “Number of Images” and “Inter-delay” were all tested and met operational expectations. The “Time Delay” function refers to the delay in seconds from the reference time for taking an image. As the name of the function indicates, the “Number of

Images” function refers to the number of images that is to be taken by the imager. The “Inter- delay” function is only applicable in the case that there is more than one image being taken, as it refers to the time delay in seconds between multiple images.

The “Time Delay” function did not work on orbit due to an inconsistency between the

CanX-2 software—CANOE (Canadian Advanced Nanosatellite Operating Environment)—and the ground software—NICE (Nanosatellite Interface and Control Environment). In NICE, the time delay referred to the number of seconds after J2000 (that is, the number of seconds after Chapter 3. CanX-7 Deployment Detection Subsystem 19

00:00:00 in year 2000) and not the time after the experiment was scheduled, as was expected.

To resolve this issue, a new version of NICE was uploaded and appeared to be functioning as expected.

3.5.2 Colour Imager Experiments

It was found that daytime images taken with minimum exposure settings were overexposed and night time images taken with maximum exposure settings were underexposed. To improve exposure quality, further images were taken during the passage of CanX-2 through penumbra with imager at nadir pointing. An STK (Satellite Tool Kit) simulation of CanX-2 helped determine the times at which the satellite would be within the penumbra. Two nadir pointing images were taken with a 2,007 s inter-delay (time between the ﬁrst and the second image) to capture an image when the satellite crossed the terminator (i.e., when CanX-2 went into and emerged from eclipse during orbit). These images proved insuﬃcient to draw any conclusions.

Therefore, 10 serial images (160 × 128 px) were taken before, during and after penumbra, in order to determine favourable lighting conditions. It was seen that the images taken during penumbra were still under-saturated. Therefore, during the next experiment, four sequential, larger (520 × 152 px) images were captured leaving eclipse, starting two minutes after penumbra with one minute inter-delay. Two of these images were under-saturated, while the other two had some distinguishable cloud patterns; one of which is shown in Figure 3.4.

3.5.3 Imaging Campaign Results

A colourizing software, “ColourizeIt,” was created by the CanX-2 co-operator, Paul Choi, to

filter the taken images, as shown in Figure 3.4. This figure displays the raw image as well as several filtered versions building up to a recognizable cloud pattern. For another experiment, a different attitude was used on CanX-2 in order to capture an image of the Earth’s horizon. Dur- ing this experiment, the image was not as sensitive to timing as much as satellite orientation.

The imager was pointed 30◦ below anti-velocity so as to capture earth, space and a gradient, indicating the atmosphere. At either extreme, the earth was expected to be completely saturated and space was expected to be under-saturated. The resulting image after post-processing with ColourizeIt software is shown in Figure 3.5. Chapter 3. CanX-7 Deployment Detection Subsystem 20

Figure 3.4: Image taken two minutes after penumbra by CanX-2 colour imager

Figure 3.5: An image of Earth’s horizon taken by CanX-2 colour imager

3.5.4 Conclusion

Even though the CanX-2 colour imager had manual control over exposure time, the resulting pictures were extremely poor and needed heavy post-processing and specific orbital conditions for acceptable quality. Thus, for the potential imager on CanX-7, a camera filter is strongly recommended for an additional requirement of the deployment detection subsystem. This is especially critical as drag sail images were intended to be captured during morning passes for full satellite illumination, and the production of washed-out images—as during CanX-2 imaging experiments—would be wholly insufficient for deployment determination. The author’s CanX- Chapter 3. CanX-7 Deployment Detection Subsystem 21

2 imaging campaign experience has therefore reinforced the belief that while an automatic exposure time would be acceptable for a potential CanX-7 imager, a ﬁlter to block out sunlight is strongly recommended as part of the design.

3.6 Current Status

After PDR, the volume allocation for the imager on the main tray was no longer available due to sequence of events that will be discussed in Section 5.3. The result was to discontinue the deployment detection subsystem (DDS) as there was an alternate method of imaging the drag sail. Therefore, no further work was done on the DDS by the author since PDR. However, as the design progressed, the alternate method of imaging was no longer available and as a result, the DDS is part of the the CanX-7 satellite again. Chapter 4

CanX-7 Preliminary System Development

At the beginning of the author’s thesis, CanX-7 was in its earliest stages of design and only two students (including the author) were involved with the mission. The second student—

Barbara Shmuel—was working on the payload while the author focused on the bus design of the spacecraft. The CanX-7 system was in a state of flux at this time because much uncertainty remained regarding its funding situation. As a result, multiple designs were carried forward in parallel. Looking at the system as a whole was very important during this stage because this was the only way to evaluate different design options for CanX-7. System level design is significant as from it all subsystem requirements are derived. This level of design involves making trades at the system level to achieve the mission goal, as well as creating and updating budgets to ensure a feasible design. This chapter describes the evolution of the preliminary system design which occurred throughout the majority of the author’s first year of Master’s work.

4.1 Driving System Requirements

The starting point of any design is the list of system requirements, which dictates what the system as a whole should accomplish. As such, the starting point for the design of CanX-7 was the “CanX-7 Mission and System Requirements” [18]. These requirements shaped the

22 Chapter 4. CanX-7 Preliminary System Development 23 preliminary design of CanX-7 and are provided in Table 4.1.

Table 4.1: Preliminary mission and system requirements No. Requirement Description Mission Requirements 1.1 Mission of CanX-7 is to demonstrate de-orbiting of a 15 kg spacecraft (spacecraft mass) from 800 km (initial orbit) within 25 years (spacecraft lifetime) through the use of a passive drag device. 1.11 A drag device (or a combination of devices) with an area of 2 m2 shall be implemented. The area requirement is based on the STK lifetime analysis using the specified initial orbit and spacecraft mass. 1.12 The drag device(s) shall be deployed by ground command(s). 1.13 If a shorter orbital lifetime (less than 25 years) is desired, it shall be achieved by a lower spacecraft mass and/or an initial orbit with a lower altitude. 1.2 The spacecraft shall aerodynamically stabilize. 1.21 The “shuttlecock effect” to aero stabilize may be be achieved through the proper Cp/Cm relationship. 1.3 A sensor(s) shall be incorporated to positively confirm the deployment of the de-orbit device (or the combination of devices). 1.31 The sensor shall be capable of determining the percentage of deployment to within ±10% (10%, 20%, . . . , 90%, 100%). 1.32 The sensor shall be capable of recording the percentage of deployment during the deployment process to within ±10%. 1.4 A sensor(s) shall be incorporated to confirm the orientation of the de-orbit device and the velocity vector. 1.41 The sensor shall be capable of determining the error in alignment between the de-orbit device normal and the velocity vector to ±5◦ or better. 1.42 The sensor shall be capable of recording the orientation between the de-orbit device and the velocity vector over a period of time. 1.5 The spacecraft shall be designed to follow either the CanX-2 bus design or the Generic Nanosatellite Bus (GNB) design. 1.51 The CanX-2 Bus: 10 × 10 × 34.5 cm form factor. 1.52 The Generic Nanosatellite Bus: 20 × 20 × 20 cm form factor, with additional appendages. 1.53 The spacecraft shall incorporate GNB components for the bus components. 1.6 The complete functionality of the spacecraft shall be testable and verifiable in 1 g, standard laboratory environment, when the spacecraft is properly supported, and without requiring any additional specialized ground support equipment to aid deployment of drag sails. System Requirements 2.1 The mass of CanX-7 shall be less than 3.5 kg in the case of a CanX-2 bus design. 2.2 The mass of CanX-7 shall be less than 7 kg in the case of a GNB bus design. 2.3 The thermal design of the CanX-7 spacecraft shall be passive and maintain all of the spacecraft components to within their respective operational limits. Chapter 4. CanX-7 Preliminary System Development 24

4.2 SFL Spacecraft Platforms

One of the goals that was very important in this early stage was making the payload compatible across the three main spacecraft platforms at SFL: CanX-2 class (triple CubeSat, or “3U”

CubeSat), Generic Nanosatellite Bus (GNB) class and Nanosatellite for Earth Monitoring and

Observation (NEMO) class (Figure 4.1). This would ensure that the drag sail design flown specifically on CanX-7 could use its earned flight heritage on future missions regardless of bus. Through CanX-7 demonstration, the technology would be qualified to instill confidence and eliminate the risk factor associated with an unproven design. Ideally, there would be no or minimal non-recurring engineering involved with the de-orbiting technology when used on future missions because the drag sail design would need little or no further modification to accommodate integration on other SFL-supported platforms.

(a) 3U platform (b) GNB platform (c) NEMO platform

Figure 4.1: Three main SFL spacecraft platforms

Triple CubeSat

The triple CubeSat (3U) platform design at SFL is based on the CanX-2 satellite which launched in April 2008. This is the only SFL mission that has used a 3U form factor to date. A solid model of it can be seen in Figure 4.1(a). It has dimensions of 100 × 100 × 340.5 mm (3U form factor) and employs older on-board technology. More information about the 3U platform may be found in Chapter 5. Chapter 4. CanX-7 Preliminary System Development 25

Generic Nanosatellite Bus

The next generation of platform and successor to the 3U form factor is the Generic Nanosatellite

Bus (GNB) platform. It was conceived at SFL to enable low cost, responsive missions with the

flexibility of using the same bus with multiple mission-specific payloads. The GNB was initially designed for CanX-4/CanX-5—a formation-flying demonstration mission—as well as BRITE

(BRIght-star Target Explorer)—a mission to examine the variability of some of the brightest stars in the sky by making photometric observations [6]. However, this platform has been extended to other missions due to its versatility. Two such examples include AISSat-1—a ship- monitoring mission which successfully launched in July 2010—as well Antarctic Broadband, whose prime objective is to demonstrate on-orbit broadband communication between research stations in the Antarctic and gateway stations on the mainland.

The GNB is a 20 × 20 × 20 cm satellite (Figure 4.1(b)). It is composed of two main trays to which the majority of components are mounted, and six aluminum panels that form the shell of the spacecraft. The GNB has a full suite of advanced capabilities which include:

• A power system with peak power tracking capabilities;

• A full 3-axis attitude determination and control system;

• An on-board computer processing system comprised of a House Keeping Computer (HKC),

an Attitude Determination and Control Computer (ADCC) and a Payload On-Board

Computer (POBC) which is a dedicated computer for payload activities;

• A full duplex communication system; and

• A payload bay in between the two main trays approximately 8 × 13 × 17 cm.

Nanosatellite for Earth Monitoring and Observation

The Nanosatellite for Earth Monitoring and Observation (NEMO) platform was derived from the GNB design for larger spacecraft. The main structure of the NEMO bus has an outer envelope of 20 × 20 × 44 cm and a mass of approximately 15 kg. NEMO shares many of the

GNB subsystems including the communication subsystem, attitude determination and control subsystem and on-board computer processing system. The power system also derives its heritage from GNB but was augmented to deliver higher power. To this end, NEMO has a large Chapter 4. CanX-7 Preliminary System Development 26

58 × 59 cm solar array attached to one side of the main structure, as shown in Figure 4.1(c).

The main structure houses all the spacecraft avionics and payload. Currently, there are two missions which utilize the NEMO bus: NEMO-AM (Aerosol Monitoring) and NEMO-HD (High

Deﬁnition).

4.3 Initial Design Options

The initial focus in CanX-7 system development was on potential bus designs and a comparison of them using various criteria. Two diﬀerent general layouts were considered for use in the

CanX-7 design—3U and GNB—based on Requirement 1.5 listed in Table 4.1: The spacecraft shall be designed to follow either the CanX-2 bus design or the GNB bus design. From this requirement, the following bus designs were initially considered for CanX-7:

1. GNB: Utilizes the GNB platform—a 20 × 20 × 20 cm form factor—with GNB avionics.

2. Hybrid: Utilizes the combination of both the 3U CubeSat and GNB platforms—a

10×10× 34.5 cm form factor with GNB avionics.

3. 3U: Utilizes the 3U CubeSat platform—a 10 × 10 × 34.5 cm form factor—using CanX-2

avionics with re-design of obsolete parts.

The listed designs will be referred to in the remainder of this chapter as Designs 1, 2 and

3, respectively. For a system level evaluation, these candidate CanX-7 designs were compared in terms of technical feasibility and development cost to select a bus. The technical feasibility of the design encompassed the following criteria: ease of manufacturability, ﬂight heritage, volume margin, mass margin and energy margin—all of which will be discussed in more detail in Section 4.4 below. The designs were also analyzed from a cost perspective, which is presented in Section 4.5.

4.4 Technical Feasibility Evaluation

For design evaluation, technical feasibility encompasses engineering factors without explicitly considering cost. These factors do not all influence the technical feasibility of a design in the same way and thus were assigned appropriate weights. In order to compare the different designs using various criteria, the raw scores assigned to each were first normalized. To normalize, the Chapter 4. CanX-7 Preliminary System Development 27

Table 4.2: Technical feasibility evaluation matrix Ease of Volume Mass Energy Flight Criteria Total Manufacturability Margin Margin Margin Heritage Options Weight 50% 15% 15% 15% 5% 100% Raw 1.00 1.00 1.27 2.12 1.00 Design 1 Normalized 1.00 1.00 1.00 1.00 1.00 (GNB) Weighted 0.50 0.15 0.15 0.15 0.05 1.00 Raw 1.00 0.00 0.82 0.24 0.00 Design 2 Normalized 1.00 0.00 0.65 0.11 0.00 (Hybrid) Weighted 0.50 0.00 0.10 0.02 0.00 0.61 Raw 0.00 1.00 0.94 0.93 1.00 Design 3 Normalized 0.00 1.00 0.74 0.44 1.00 (3U) Weighted 0.00 0.15 0.11 0.07 0.05 0.38 raw score of each design was divided by the greatest raw score for that category such that the greatest normalized score was one. Next, the weighted score for each design was calculated by multiplying the normalized scores with their respective weights. The following subsections will explain the metrics used and weights assigned for each criterion, as well as the raw score that was assigned to each design. Finally a total number is calculated for each of the three designs by adding all the weighted scores. This total is implicitly normalized due to the normalizing of individual criteria prior to multiplying by weights. A higher total score indicates a more technically feasible design. Table 4.2 summarizes the evaluation results.

4.4.1 Ease of Manufacturability

Ease of manufacturablity refers to the practicality of being able to physically construct the design. This was assigned a weight of 50% as the author believes this is an extremely important factor (moreso than other criteria) as the end goal is to build a real satellite from the chosen design. The scoring for this criteria was dependent on whether the components needed for each design were readily available and easily obtainable. A score of one indicates that all components are available. This was the score for Designs 1 and 2, as they utilize GNB components—which are the current technology. As a result of the technical and functional obsolescence of CanX-2 components, some are no longer available. Therefore a raw score of zero was assigned to Design

3. Chapter 4. CanX-7 Preliminary System Development 28

4.4.2 Volume Margin

The SolidEdge software was used to model the three designs and aided in the volume margin evaluation. This criterion along with the other two margin criteria (mass and energy) were assigned equal weights of 15% as they are the primary aspects considered early in a design.

The volume budget helped visualize how the payload and avionics (CanX-2 or GNB) would

ﬁt into the corresponding form factor (3U or GNB) of each design. A volume margin was assessed based on the excess volume left over after all components were assembled. Since this assessment was qualitative, a score of one or zero implied that there was positive or negative margin, respectively.

Design 1: Generic Nanosatellite Bus

Design 1 was a GNB bus with GNB components and, as a result, there were no signiﬁcant problems in ﬁtting all components inside. During the early stage of development, the payloads for CanX-7 consisted of four drag sail modules, each containing a drag sail as well as two

Sinclair star trackers that were to be used as cameras for imaging the sail after deployment.

(This component design predated Deployment Detection Subsystem (DDS) development.) The drag sail modules were laid out on a plane, as seen on the left side of Figure 4.2, on the −Y face to avoid interfering with antennas during deployment. In order to leave some gap between the top cover of the drag sail module and the feet of the launch rails which will touch the XPOD—

SFL’s separation system to eject satellites into orbit—the height of each drag sail module was reduced to 0.8 cm from 1 cm. Reducing the height of the drag sail module was preferred over adjusting the GNB tray (by increasing the leg length of the launch rails) because the drag sail module was in the design stage and this decision maintains the structural heritage of the GNB tray, without introducing the need for non-recurring engineering. Therefore, the GNB tray was not modiﬁed in any way.

Initially, four star trackers were to be used—one to image each sail. However, due to limited space, this was reduced to two. The two star trackers were to face opposite directions to capture the maximum area of the sail. In order to image all four sails, each star tracker would image a portion of two sails. They were not able to be placed on the +Z and −Z faces since there was Chapter 4. CanX-7 Preliminary System Development 29 no space on the trays to insert the star tracker. There was also no space on the +Z panel for the cutout required for the star tracker lens since the magnetometer and test port were placed on that face. Therefore, the star trackers were placed on the +X and −X faces. From the right side of Figure 4.2 it may be seen that a casing was added on both +X and −X faces in order to house the star trackers, and that these casings did not interfere with the dual trays.

As all the avionics and payloads were able to ﬁt inside the 20 × 20 ×20 cm volume, Design 1 was assigned a raw score of one for the volume budget criteria.

Figure 4.2: Drag sail module mounting scheme and star tracker layout on the GNB

Design 2: Hybrid

The CanX-2 form factor is a 100 × 100 × 340.5 mm volume with the original satellite possessing a main tray on which most of its components were mounted. Since GNB components are to be used for Design 2 of CanX-7, the original main tray was modiﬁed to accommodate the

GNB components. However, this design presented a geometric problem as it used a 3U form factor with GNB components. SolidEdge modeling demonstrated that a GNB power board was physically unable to ﬁt within the 100 × 100 × 340.5 mm volume of CanX-2 even with the modiﬁed layout as its dimensions were 164 × 136 × 2 mm. This resulted in a negative volume margin and therefore a nil score in the technical feasibility evaluation for Design 2.

In addition to the avionics, the 3U form factor needed to ﬁt the four drag sail modules

(stacked on top of each other) as well as the two star trackers. To ﬁt all drag sail modules on Chapter 4. CanX-7 Preliminary System Development 30 top of each other, the legs on the launch rails of the tray required lengthening. To ﬁt the two star trackers with the rest of the components in the CanX-2 form factor, one star tracker was placed near the +Y end and the other near the −Y end with apertures on +Z and −Z. From

Figure 4.3(a), it may be seen that the star tracker near the −Y end was tilted toward +Y since its field of view without tilt is insufficient to image any part of the sail. In contrast, the star tracker near the +Y end does not need tilting since its field of view is enough to image the sail from its position on the main tray. The modified, populated (except for drag sails) main tray of Design 2 is shown in Figure 4.3(a). The deployed configuration of Design 2 is shown in Figure 4.3(b) along with the field of view of each of the star trackers. Each star tracker is able to capture a portion of two of the drag sails. Therefore, their positions and orientations are sufficient for the purpose of confirming sail deployment.

(a) Modiﬁed main tray (b) Deployed conﬁguration

Figure 4.3: Design 2: Hybrid

Design 3: Triple CubeSat (3U)

Similar to Design 1, Design 3 also did not pose any signiﬁcant problems since it had a CanX-2 form factor with mostly CanX-2 parts, and thus resulted in a score of one in the evaluation matrix. However, the re-design of the obsolete parts necessitates conﬁrmation of all components

fitting within a CanX-2 volume. For this design, even though CanX-2 components were used, the main tray had to be modified to in order to accommodate the four drag sail modules and two star trackers. Similar to Design 2, the launch rails of the tray had to be lengthened to Chapter 4. CanX-7 Preliminary System Development 31 stack the drag sail modules. The tray for Design 3 is shown in Figure 4.4(a) and the deployed configuration is given in Figure 4.4(b).

(a) Modiﬁed main tray (b) Deployed conﬁguration

Figure 4.4: Design 3: 3U

4.4.3 Mass Margin

Like the volume margin, a weight of 15% was also assigned to the third criterion: mass margin.

The mass margin was derived from the preliminary mass budget of the three designs presented in Table 4.3. The mass margin numbers were not used directly as two designs had negative margins while the third was positive. The total to target mass ratio was used instead because this captured the mass margin and provided a way to compare all three designs with positive numbers. This ratio essentially provided the mass margin in a diﬀerent form with a healthy margin corresponding to a smaller ratio. As a larger score corresponded to a better design in the evaluation matrix, the inverse of the ratio was used, resulting in a target to total mass ratio.

This corresponded to raw scores of 1.22, 0.79, and 1.06 for Designs 1, 2 and 3 respectively.

4.4.4 Energy Margin

Energy margin, like the criteria of volume and mass margins, was also assigned a weight of 15%.

Data for this metric was derived from the energy budgets created for all three designs, whose results are provided in Tables 4.4 through 4.6. Having a large energy margin was important because, at this time, CanX-7 was very early in its design stage. Chapter 4. CanX-7 Preliminary System Development 32

Table 4.3: Preliminary mass budget for three candidate CanX-7 designs Design 1: GNB Design 2: Hybrid Design 3: 3U Subsystem Mass (g) Fraction (%) Mass (g) Fraction (%) Mass (g) Fraction (%) Structural 2,569 46 1,327 31 1,327 36 Thermal Control 50 1 50 1 50 1 ADC 148 3 148 3 277 7 Power 1,186 21 1,186 21 429 12 Computer 63 1 63 1 79 2 Communications 259 5 259 6 311 8 Payloads 1,210 22 1,210 28 1,210 32 Integration 55 1 42 1 37 1 Total 5,541 100 4,286 100 3,726 100 Target 7,000 – 3,500 – 3,500 – Margin 1,459 21 -786 -22 -226 -6 Target/Total 1.22 0.79 1.06

Table 4.4: Energy budget for Design 1: GNB Safe Hold Commission Payload Sail Deploy De-orbit Consumed (W·hr) 1.63 2.03 2.08 2.21 2.13 Generated (W·hr) 5.56 5.56 5.56 5.56 5.56 Energy Margin (%) 70.6 63.5 62.5 60.2 61.7

Table 4.5: Energy budget for Design 2: Hybrid Safe Hold Commission Payload Sail Deploy De-orbit Consumed (W·hr) 1.63 2.03 2.08 2.21 2.13 Generated (W·hr) 1.86 2.25 2.25 2.25 2.25 Energy Margin (%) 12.2 9.8 7.5 1.7 5.5

Table 4.6: Energy budget for Design 3: Triple CubeSat (3U)

Safe Hold Commission Payload Sail Deploy De-orbit Consumed (W·hr) 1.32 1.54 1.59 1.72 1.64 Generated (W·hr) 1.86 2.25 2.25 2.25 2.25 Energy Margin (%) 28.8 31.5 29.2 23.4 27.2 Chapter 4. CanX-7 Preliminary System Development 33

From Tables 4.4 to 4.6, all spacecraft modes for all three designs are shown to have positive energy margins. However, certain designs had signiﬁcantly higher energy margins than others.

This early in the design phase, a 30% energy margin was usually required for typical SFL satellites. Therefore, the energy metric was the ratio of the average energy margin of all spacecraft modes to the desired 30% value. A weighted average of the modes could have been used for criterion reﬁnement, but this was not expected to change the results signiﬁcantly due to low margin spreads for each design.

All the modes in Design 1 had at least double the typically required energy margin (63.7% average) while most of the modes in Design 3 had nearly a 30% energy margin (28.0% average).

Design 2, on the other hand, was far from having energy margins of 30%. In fact, Design 2 marginally closed the energy budget (7.3% average). This is because the GNB components nearly required more power than the CanX-2 form factor was able to provide. In the order of decreasing energy margin, Designs 1, 3 and 2 scored raw values of 2.12, 0.93 and 0.24, respectively, in the technical feasibility matrix.

An altitude range of 600 to 800 km was used for orbit calculations as CanX-7 is expected to be launched within this region. The following conservative assumptions were further made in the generation of the energy budget: the satellite is in a noon-midnight orbit with an orbital period of 1.5 hours and the Earth has albedo of zero. The worst-case orbit for a satellite is a noon-midnight orbit since this corresponds to the greatest time in eclipse and therefore the lowest amount of time for power generation through solar arrays. An albedo of zero is also a worst-case scenario because it assumes no sunlight reﬂection oﬀ the Earth (which is most unlikely).

The worst-case attitude for Design 1 is when only one face is pointing towards the sun. Since it is a cube, it does not matter which face is towards the sun as all six have equal number of solar coupons and therefore will produce the same amount of power. Initially for Designs 2 and

3, which use a CanX-2 form factor, the worst-case attitude was considered with the smallest face always pointing towards the sun. This produced negative energy margins for all of the spacecraft modes. Therefore, this worst case attitude was only considered for the safe-hold mode as there is no attitude control. Furthermore, the duty cycle of certain units that could be adjusted were lowered to close the energy budget. For the rest of the modes, a more realistic Chapter 4. CanX-7 Preliminary System Development 34 case for CanX-7 attitude was considered which used a spherical average that corresponds to a random tumble.

4.4.5 Flight Heritage

The last criterion that was used in the evaluation of the technical feasibility was ﬂight heritage.

This was assigned a weighting of 5% as, compared to the previously identified factors, flight heritage was considered least important for a novel technology demonstration mission. As they have flown on previous SFL satellite missions, both the CanX-2 and GNB designs (and their corresponding components) have flight heritage. A number in this criterion corresponds to the number of satellites that have flown with the design. Therefore, Designs 1 and 3 scored one each as there is only one GNB satellite (AISSat-I) and one CanX-2 bus design (CanX-2) currently in orbit. Design 2 scored nil as the integration of a CanX-2 bus with GNB components in one system has no flight heritage.

4.4.6 Technical Feasibility Results

The technical feasibility evaluation indicated that Design 1 (GNB) would be the best option, without consideration of cost, with a “perfect” total score of one. The next ideal design would be Design 2 (Hybrid) with a score of 0.61 followed by Design 3 (Triple CubeSat or 3U) with a score of 0.38.

4.5 Cost Evaluation

A cost analysis was performed for comparison of the three potential CanX-7 bus designs as it was also an engineering factor that inﬂuenced the design decision. To determine the best design for CanX-7, “value” was used as the ﬁgure of merit rather than the development cost.

The value number indicates the worth of each design and the signiﬁcance each has to SFL.

The value of each design was calculated as the ratio of its beneﬁt to its development cost. The

ﬁrst part of this section discusses the CanX-7 development cost and its method of calculation.

Following this is a description of the beneﬁt calculation, as well as the resulting value of each potential CanX-7 design. Chapter 4. CanX-7 Preliminary System Development 35

4.5.1 Development Cost

The categories taken into account to calculate the cost of CanX-7 were non-recurring engineering, materials, unit-level assembly, integration and testing (AIT), ﬂatsat integration cost and launch. This is a fairly complete list of costs that captures the development process of a satellite from design to launch. The labour rate that was used in some of the calculations assumed a cost of US$100,000 per annum per worker. This rate is arbitrary, and does not represent actual labour cost, but is used as a basis of comparison only.

Non-Recurring Engineering Cost

The cost analysis presented here is intended to focus on diﬀerentiating the design options only, not on absolute costs. In this case, non-recurring engineering (NRE) cost refers only to the bus portion, as the NRE cost of the payload will be equal for all designs. The considered NRE cost includes the design of new components as well as the cost in qualifying these components.

For Design 1, there is no cost involved with NRE as the GNB is an established platform.

A GNB requires no further modification as all components needed for CanX-7 fit within the allocated region. Design 2 involves fitting GNB components into a CanX-2 form factor. From solid modeling, all components except the GNB power board are able to fit within a CanX-2 form factor. Therefore, NRE costs for Design 2 are attributed to designing and qualifying a

GNB power board to ﬁt within a 3U form factor as well as modiﬁcation of the main tray to accommodate all GNB components in a 100 × 100 × 340.5 mm volume.

Design 3 involves using a CanX-2 form factor with mostly CanX-2 components with the replacement of obsolete parts. The FPGA chip on the power board is one of the main obsolete part from CanX-2. In building electrical, electronic or electromechanical parts for SFL satellites, military standard MIL-STD-975 is typically followed [7]. The NRE cost of designing a CanX-2 power board adhering to MIL-STD-975 will be different than designing a CanX-2 power board with exceptions to MIL-STD-975, as was the case during the design of the original CanX-2 satellite. Since the difference in cost is significant between the two cases, the Design 3 option is split into two designs (3a and 3b). The power board in Design 3a is designed with exceptions to the MIL-STD-975, and the power board in Design 3b is designed adhering to the MIL-STD-975. Chapter 4. CanX-7 Preliminary System Development 36

Table 4.7: Non-recurring engineering cost breakdown by labour time of activity Design Design Design Design Task 1 2 3a 3b Design of Schematic (wks) 0 4 1 4 Power PCB updates (wks) 0 4 2 4 Board FPGA programming (wks) 0 3 1 1.5 Assembly of board (wks) 0 1.5 1.5 1.5 Qualiﬁcation Full functional test (wks) 0 1 1 1 of Power Radiation test (wks) 0 0.5 0.5 0.5 Board Radiation checkout (wks) 0 1 1 1 TVAC Test (wks) 0 1.5 1.5 1.5 Total time (wks) 0 16.5 9.5 15 Total time with contingency (wks) 0 24.75 14.25 22.5 Labour rate ($/wk) 1,923 1,923 1,923 1,923 Total NRE Cost ($) 0 47,596 27,404 43,269

The breakdown of the NRE cost for each design is presented in Table 4.7. Labour time was estimated based on SFL engineering expertise. A margin of 50% contingency is added to each labour total. The total NRE cost is calculated by multiplying the total time with contingency by the labour rate. Designs 3a and 3b are carried forward in the rest of the cost categories for completeness.

Raw Material Cost Diﬀerential

Material cost refers to the cost of components and hardware required to build the diﬀerent units. Only components that are diﬀerent between designs are used for this calculation. From

Table 4.8, the main diﬀerence in the material cost of a GNB class (Design 1) compared to a CanX-2 class (Designs 2, 3a, 3b) satellite is due to the number of solar cells required, as the price of each is very signiﬁcant in comparison to other components. A GNB, with greater surface area than CanX-2, supports six solar cells on each face (36 total). CanX-2 form factors can only support a total of 18 solar cells. Chapter 4. CanX-7 Preliminary System Development 37

Table 4.8: Raw diﬀerential material cost break down Design 1 Design 2 Design 3a Design 3b Unit Unit Cost ($) No. Cost ($) No. Cost ($) No. Cost ($) No. Cost ($) BCDR 150 1 150 1 150 0 0 0 0 Battery 150 2 300 2 300 1 150 1 150 Solar Cell 1,475 36 53,100 18 26,550 18 26,550 18 26,550 Tray (GNB) 2,300 2 4,600 0 0 0 0 0 0 Panel (20×20 cm) 883 6 5,298 0 0 0 0 0 0 Interior Tray (CanX-2) 1,500 0 0 1 1,500 1 1,500 1 1,500 Panel (CanX-2) 883 0 0 6 5,298 6 5,298 6 5,298 Total Materials Cost ($) 63,448 33,798 33,498 33,498

Unit-level Assembly, Integration and Testing Cost Diﬀerential

Unit-level assembly, integration and test (AIT) costs refers to the cost of building and testing each unit (component) on the bus. The units included in Table 4.9 are those built in-house that are not in common among the designs. From this table, it may be seen that Designs 1 and

2 have the same unit-level diﬀerential AIT cost since both use GNB components. Similarly,

Designs 3a and 3b also share the same unit-level diﬀerential AIT cost since both use CanX-2 components. The AIT costs for each of the individual units were calculated in detail and may be found in [8].

Table 4.9: Assembly, integration and testing diﬀerential cost break down Design 1 Design 2 Design 3a Design 3b Units AIT Cost ($) AIT Cost ($) AIT Cost ($) AIT Cost ($) BCDR 4,466 4,466 0 0 Total Diﬀerential AIT Cost 4,466 4,466 0 0

Flat Satellite Integration Cost Diﬀerential

A flat satellite (or “flatsat”) is a two dimensional integration of a satellite that has all components laid out and connected to allow for testing before building the assembled design for flight.

The ﬂatsat integration cost is proportional to build-time, which was estimated for each design based on experience with past CanX-2 and GNB missions. As seen in Table 4.10, the ﬂatsat of Design 2 would take longest to build as this design is a hybrid of both GNB and CanX-2 classes, and has never been built before. Chapter 4. CanX-7 Preliminary System Development 38

Table 4.10: Flat satellite (ﬂat-sat) cost break down

Design 1 Design 2 Design 3a Design 3b Time (wks) 3 10 4 6 Labour Rate ($/wk) 1,923 1,923 1,923 1,923 Total Flat-Sat Cost ($) 5,769 19,231 7,692 11,538

System-level Assembly, Integration and Test Cost

System-level assembly, integration and test (AIT) costs refer to the cost of assembling the entire spacecraft after all the component units have been built and tested. This cost also includes black box testing. The time it takes to build and test either a GNB class or a CanX-2 class spacecraft is the approximately the same and therefore all four potential CanX-7 designs will have the same system-level AIT cost. This cost is therefore not included in the cost comparison.

Launch Cost

The launch cost of each design is proportional to the weight of the satellite itself and its corresponding separation system, XPOD, which attaches to the launch vehicle. Depending upon the launch provider, the cost per kilogram could vary. An arbitrary number for comparison purposes of e20,000 (US$27,023) per kg was used. Design 1 has signiﬁcantly larger launch cost than the other three designs since GNB is a bigger and heavier satellite with a correspondingly larger XPOD. The comparative launch cost of each design is given in Table 4.11.

Total Diﬀerential Development Cost and Launch Comparison

The total differential development cost of each of the potential CanX-7 designs are presented in Table 4.12. It may be seen that a GNB design is the most costly and CanX-2 built with exceptions to the military standard (Design 3a) is the least costly. Development costs shown here are differential only and do not reflect the base recurring cost of each satellite platform.

However, the development cost alone is not suﬃcient to determine the best design for CanX-7.

The next section (4.5.2) will address the value of building CanX-7 in each of these designs. Chapter 4. CanX-7 Preliminary System Development 39

Table 4.11: Launch cost break down Design 1 Design 2 Design 3a Design 3b Subsystem Mass (g) Mass (g) Mass (g) Mass (g) Structural 2,569 1,327 1,327 1,327 Thermal Control 50 50 50 50 ADC 148 148 277 277 Power 1,186 1,186 429 429 Computer 63 63 79 79 Communications 259 259 311 311 Payloads 1,210 1,210 1,210 1,210 Integration 55 42 37 37 XPOD 9,000 3,000 3,000 3,000 Total Mass (g) 14,540 7,285 6,720 6,720 Cost/Mass ($/kg) 27,023 27,023 27,023 27,023 Total Launch Cost ($) 392,921 196,866 181,598 181,598

Table 4.12: Total differential development cost breakdown Type of Cost Design 1 Design 2 Design 3a Design 3b Design NRE ($) 0 47,596 27,404 43,269 Differential Material ($) 63,448 33,798 33,498 33,498 Differential Unit-level AIT ($) 4,466 4,466 0 0 Differential Flatsat Integration ($) 5,769 19,231 7,692 11,538 Launch ($) (arbitrary rate) 392,921 196,866 181,598 181,598 Total Cost ($) 466,604 301,957 250,192 269,903

4.5.2 Beneﬁt and Value

In order to obtain the value of each potential CanX-7 design, the benefit of each design had to be quantified. In terms of benefit, SFL managers estimate that the GNB brings approximately 18 times the benefit to clients as the 3U (CanX-2) bus. This is based on the number of applications and interest from clients. The satellites of SFL were used as a representative sample of missions in Canada; this method is slightly qualitative but not unreasonable as SFL is one of the leaders in its field. Developing CanX-7 as a GNB is significantly more beneficial than developing it as a CanX-2. The value of each design was calculated using the formula below and the result is presented in Table 4.13.

Beneﬁt Value = Development Cost Chapter 4. CanX-7 Preliminary System Development 40

Table 4.13: Value of CanX-7 Design 1 Design 2 Design 3a Design 3b Beneﬁt 18 1 1 1 Development Cost ($) 466,604 301,957 250,192 269,903 Normalized Value 1 0.085 0.10 0.095

4.6 Preliminary Design Decision

Developing CanX-7 as a GNB would be of signiﬁcantly greater value than 3U, as found in

Subsection 4.5.2, with a value number of an order of magnitude greater than competing designs.

Also, the technical feasibility matrix (Table 4.2) indicates that the GNB design for CanX-7 would be the most feasible solution. Weighing against these advantages is the development cost, which is highest for Design 1.

When working on a real-life engineering problem such as the CanX-7 mission, the available budget plays a signiﬁcant role and drives certain decisions. The funding situation at this point in the design process did not allow for a GNB class mission; therefore, the baseline was a

3U form factor with the option of porting the mission to a GNB. The system level trade was technical feasibility and value for development cost. The limited budget drove the decision to pursue the smaller form factor, as the manufacturing, testing and the launch would be more cost-eﬀective than for a larger form factor. The 3U form factor would also limit the number of components the satellite could carry, which in turn would reduce the overall satellite cost.

Thus CanX-7 had to be Design 2, 3a or 3b.

Between Designs 2 and 3 (a and b), the technical feasibility evaluation in Table 4.2 indicated

Design 2 was better. And since the option of growing to a GNB class mission was still not completely dismissed, Design 2 (Hybrid) seemed favourable as it uses GNB avionics. Although the hybrid design itself would be more diﬃcult to develop than a 3U (the structure required signiﬁcant design work to accommodate components that were initially not present), it was more practical to build, since all the components could be purchased without an extensive lead time. Therefore, based on the preliminary system analysis performed for CanX-7, the decision was to use the hybrid design for CanX-7 spacecraft. Chapter 5

CanX-7 Structural Subsystem

Once the baseline design was chosen and as the CanX-7 team grew, the focus of the author’s work shifted from system to subsystem design—in particular that of the CanX-7 structure.

This was the most logical subsystem to work on as much of the work done at the system level was closely tied to the structural aspect of the satellite.

The main purpose of the structural subsystem is to provide structural support for other subsystems and payloads during integration, testing, transport and launch. Due to the nature of the design process, the structural subsystem has undergone multiple design changes over the course of this Master’s thesis. Discussed in this chapter are two iterations of the structural design: an initial design presented at the preliminary design review (PDR) as well as the current design.

SolidEdge software has aided the structural design of CanX-7. This software was mainly used for design of primary structural components as well as determining placement of avionics.

It was decided that wiring harness design could be best determined by building a mock-up of the satellite rather than trying to visualize it using the CAD model. Therefore, after a thorough review of the CAD model by an SFL staﬀ engineer, Cordell Grant, the CanX-7 structure was sent out for mock-up manufacturing. In parallel with building the mock-up satellite, structural analysis was performed using NX software. A ﬁnite element model (FEM) was created and structural simulations were performed for modal analysis and for evaluation of the satellite to handle launch loads. This chapter summarizes the design evolution of the structural subsystem, the process of manufacturing and building the mock-up satellite, the results from structural

41 Chapter 5. CanX-7 Structural Subsystem 42 analysis and the future work remaining for the CanX-7 spacecraft.

5.1 Structural Requirements

The system level requirements that most heavily impact the CanX-7 bus mechanical design are listed below. These are from the System Requirements Review (SRR) meeting that took place in

August 2011, or approximately halfway through the author’s Master’s thesis. The requirements listed in Section 4.1 were used as the baseline for the SRR. From this point onwards, the focus of the author has been on the structural subsystem. A complete list of system requirements can be found in [11].

SYS-31 The mission should use CanX-2 and/or GNB heritage components to the extent prac-

tical.

SYS-32 The spacecraft dimensions, including appendages, shall be compatible with a qualiﬁed

XPOD deployment system.

SYS-33 The CanX-7 spacecraft should not require any modiﬁcations to the selected standard

XPOD design.

SYS-46 If a triple CubeSat form factor is used, the satellite mass shall not exceed 3.5 kg.

SYS-77 The spacecraft design should permit easy partial assembly and disassembly that may

be done using standard laboratory tools.

SYS-85 In its ﬂight conﬁguration the satellite shall be subjected to an acceptance-level vibra-

tion test at levels speciﬁed by the launch provider.

SYS-86 The spacecraft shall be tested in a thermal vacuum environment at the extremes of

its operational temperature range.

5.2 CanX-7 Preliminary Structural Design

A preliminary structure was designed to meet all the requirements stated in Section 5.1. This section describes the CanX-7 structural design as presented at the CanX-7 PDR (Preliminary

Design Review) which took place in February 2012. Subsection 5.2.1 explains the structural concept, and is followed by the structural modiﬁcations made to the CanX-2 spacecraft in

Subsection 5.2.2. Finally, the spacecraft layout during PDR is described in Subsection 5.2.3. Chapter 5. CanX-7 Structural Subsystem 43

5.2.1 Structural Concept

The ﬁrst aspect of the preliminary structural design is the structural concept. Three main factors make the CanX-7 structural concept unique: CanX-2 form factor, GNB avionics and

CanX-7 mission speciﬁc payloads. These three factors are described below.

CanX-2 Form Factor

The preliminary structural design of CanX-7 was largely based on the structural concept of the

CanX-2 spacecraft. The CanX-2 structure is primarily composed of a main tray to which the heavier components are mounted to and six panels all machined from aluminum. The CanX-7 spacecraft consists of one main tray and one main panel (−Z panel) to which the majority of the components are mounted, and four other panels (±X, ±Y ) to enclose the structure and help support a few lighter components.

(a) CanX-2 exploded view (b) CanX-7 exploded view

Figure 5.1: Structural concept of both CanX-2 and CanX-7 satellites

GNB Avionics

From the baseline design, CanX-7 was to use GNB technology. However, due to the nature of the mission, the entire suite of GNB components would not be used as CanX-7 does not require it to meet its mission requirements. At this time, it was to only carry the following

GNB components on-board:

• House Keeping Computer (HKC)

• Power board Chapter 5. CanX-7 Structural Subsystem 44

• Battery Charge/Discharge Regulator (BCDR)

• Battery

• Magnetometer

• Magnetorquer (×3)

• S-band radio

• UHF radio

CanX-7 Payloads

In addition to the bus components, CanX-7 was to also carry three payloads. The primary payload for CanX-7 are four drag sails, as the purpose of the mission is to demonstrate a novel de-orbiting technology. At this time, each drag sail module had a wedge-shaped footprint with a 3 cm height, as shown in Figure 5.2(a). The four drag sail modules will be arranged in the conﬁguration shown in Figure 5.2(b), and will be stacked on top of the rest of the CanX-

7 bus. During PDR, the CanX-7 secondary payload provided by COM DEV had maximum volume of 100 × 90 × 60 mm and maximum mass of 1 kg. An imager was also to be part of the CanX-7 payload. The imager was to conﬁrm deployment by capturing an image of the deployed sail, which would satisfy the following system level requirements: “The system shall gather engineering telemetry that can be used to detect or infer sail deployment,” and “The system should be capable of determining the percentage of sail deployment to within ±10%.”

(a) Drag sail design at PDR (b) Arrangement of the drag sail modules on CanX-7

Figure 5.2: Drag sail design and stacked conﬁguration Chapter 5. CanX-7 Structural Subsystem 45

5.2.2 Structural Modiﬁcation

In looking at the structural design as a whole, the goal was to design the CanX-7 structural components (main tray, −Z, ±X, ±Y panels) based on the CanX-2 satellite to accommodate the GNB components as well the mission specific payloads (drag sail modules, imager, and secondary payload). Toward this end, both the CanX-2 structural components as well as the structural aspects of certain GNB avionics required modification, as the CanX-2 bus was not originally designed to carry GNB technology. These modifications—made to the main tray, solar panels and GNB components—are discussed in this subsection, and comprise the second aspect of the CanX-7 preliminary structural design.

Main Tray

The ﬁrst major modiﬁcation of the CanX-2 main tray was the feet of the launch rails (not the entire tray). Each foot was lengthened by 60 mm from its original height of only 4.5 mm in order to accommodate the four drag sail modules to be stacked on top of the bus (Figure 5.3(a)).

Secondly, an allocated area of 100 × 90 mm was set aside beneath the elongated launch rails for the secondary COM DEV payload. There was no exact mounting location for this payload since its details remained uncertain.

Aside from the primary and secondary payloads, the main tray also serves as a structural support for the bus components. The largest GNB component housed by the tray is the S- band radio. To conserve mass and volume, the S-band enclosure was built into the main tray

(Figure 5.3(a). For the same reason, the +Z solar panel that originally attached to the back of

CanX-2 tray, as shown in Figure 5.1(a), was also built into the CanX-7 tray itself (Figure 5.3(b)).

As a result, there are two cutouts on the tray above the feet of the launch rails (for the coaxial cables to run from the UHF Receiver to the antennas), whereas on the CanX-2 structure, these cutouts were on the +Z solar panel (and not on the tray). Beneath the S-band enclosure was a cutout of 23.5 × 56 × 28 mm for the third payload: the imager. This allocation was as close to the −Y face as possible such that it was the greatest distance away from the drag sail modules.

This position aﬀords the greatest possible ﬁeld of view for the imager, thereby maximizing sail area that can be captured by an image. Chapter 5. CanX-7 Structural Subsystem 46

(a) Component allocation (b) Built-in solar panel

Figure 5.3: CanX-2 and preliminary CanX-7 main tray

The launch rails on the main tray were modified with cutouts such that the ±X panels would attach on top of them. The ±X solar panel mounting locations on the main tray are also different than they were on the original CanX-2 tray due to the internal layout of the satellite at PDR. On CanX-2, the ±X panels were attached to the main tray through stiffeners, whereas in the preliminary CanX-7 design, they attach to the main tray orthogonally as shown in Figures 5.4(a) and 5.4(b). The risk of screws jamming within the XPOD is a design flaw in the preliminary design of the main tray and it is later addressed in the Section 5.3 on Design

Evolution.

(a) Main tray +X mounting holes (b) +X solar panel attachment

Figure 5.4: Modiﬁed main tray and +X solar panel Chapter 5. CanX-7 Structural Subsystem 47

Solar Panels

In addition to the tray, there were ﬁve magnesium walls or solar panels that served as the structural shell for CanX-7. In the CanX-2 3U platform, there were six panels (Figure 5.1(a)).

As stated above, for CanX-7, the +Z panel which originally attached to the main tray was now built into the CanX-7 tray itself; hence, CanX-7 has one fewer panel than the CanX-2 design.

The −Z solar panel with launch rails built-in (Figure 5.5(a)) was also modiﬁed to accommodate the drag sails, which corresponded to a 60 mm extension of the feet of the launch rails (while keeping the same entire length of the −Z panel). This −Z panel also has a modiﬁed mounting location for the GNB HKC, power board and the Z magnetorquer. As the alignment of the −Z solar panel is normal to the main tray (Figure 5.5(b)), the +X solar panel mounting location on the −Z panel is identical to that of the main tray. The +X panel mounting holes and its attachment to the −Z panel are shown in Figures 5.6(a) and 5.6(b).

(a) CanX-2 and CanX-7 -Z panel (b) Main tray and −Z panel

Figure 5.5: Modiﬁed −Z panel and arrangement with main tray

The dimensions of the ±X solar panels were modiﬁed from the original panels with a 60 mm decrease along the long (Y ) axis to accommodate the drag sail modules. In order to attach orthogonally to the main tray and the −Z panel, the ±X solar panels were lengthened to Chapter 5. CanX-7 Structural Subsystem 48

(a) +X panel mounting holes on −Z panel (b) +X panel attachment to −Z panel

Figure 5.6: Modiﬁed −Z panel and +X solar panel

100 mm along the Z axis (Figures 5.7(a) and 5.7(b)) such that they spanned the width of the satellite. As a result, the launch rails on the main tray and the −Z panel have cutouts for

ﬁtting of the ±X panels. The ±X panels also have cross braces built into them to provide rigidity to the thin sheets. The stiﬀeners do not extend the entire width of the panel because this would interfere with the launch rails of the main tray and the −Z panel; thus, they do not serve as mounting points to receive screws from the tray or −Z panel (as they did in CanX-2).

The stiﬀening cross braces on the modiﬁed +X and −X solar panels are not identical to each other due to the internal layout of the satellite. The S-band patch antenna cutouts on both panels were increased by 120 mm along the long axis due to the S-band Tx connectors located near the center of the satellite.

The two ±Y solar panels are very similar to the original design used in CanX-2. For CanX-7, the only diﬀerence is that these panels have four mounting positions instead of six (Figure 5.8) as the solar coupons span almost the entire width of the satellite, thereby leaving insuﬃcient room for mounting points on the two edges at both ±X extremes.

GNB Components

There still existed some uncertainty associated with the component suite as well as the design of certain avionics and payload at PDR. In these cases, reasonable assumptions were made and the design was carried forward. Of the components that were on-board the CanX-7 spacecraft, Chapter 5. CanX-7 Structural Subsystem 49

(a) CanX-2 and CanX-7 +X panel (b) CanX-2 and CanX-7 −X panel

Figure 5.7: Modiﬁed X solar panel

Figure 5.8: Modiﬁed −Y solar panel certain structural aspects of the following components were modiﬁed such that they could be attached to the main tray or −Z panel:

• S-band radio

• UHF radio

• Power board

• Magnetometer

• Magnetorquer Chapter 5. CanX-7 Structural Subsystem 50

S-band Radio

The S-band radio design at PDR consisted of an S-band enclosure, S-band transmitter (Tx),

S-band radio cover and two S-band patch antennas. In a GNB satellite, the S-band and UHF enclosures are built into the Z− tray (Figure 5.9(a)) and share a common radio cover which does not fit within the dimensions of the 3U main tray. Since the S-band radio enclosure was built within the 3U main tray (as done on the GNB tray), the S-band transmitter is able to fit within it. However, the radio cover was modified. It was shortened to 86 mm along the width of the S-band enclosure and made specifically to cover the S-band radio. In addition to being shortened, the cover was also modified in thickness. The process of modifying the GNB radio cover to specifically cover the S-band is shown in Figure 5.9(b).

(a) GNB Z− Tray (b) Process of creating the S-band radio cover from GNB radio cover

Figure 5.9: GNB tray and radio cover

UHF Radio

The UHF radio design at PDR consisted of a UHF casing, UHF Receiver, UHF radio cover, and four UHF monopole antennas. As stated above, for a GNB, the UHF casing is built into the tray (Figure 5.9(a)). This was not possible for the main tray of CanX-7 due to space limitations. Therefore, a separate UHF casing was to be machined (Figure 5.10(a)), whose enclosure is identical to that on the GNB tray. Thus, no modiﬁcations would be necessary to the UHF Receiver, which will ﬁt within the machined casing. During assembly of the satellite, the UHF radio was to be stacked on top of the S-band radio and also mounted to the main tray. For mounting purposes, additional features external to the UHF casing were required

(Figure 5.10(a)). This was to avoid interference with the internals of the casing where the UHF Chapter 5. CanX-7 Structural Subsystem 51

Receiver was housed. In addition, a radio cover speciﬁc for the UHF radio was created because the GNB radio cover was too large to ﬁt within the 3U form factor. Therefore, the GNB radio cover width was shortened to match the width of the UHF casing such that it would be able to cover the UHF Receiver. Protrusions were added to the UHF cover such that the BCDR could be mounted on top of the UHF radio. The UHF cover is shown in Figure 5.10(a).

The last aspect of the UHF radio that needed modiﬁcation was the pre-deployed monopoles, because the XPOD-Triple—the separation system for an SFL 3U satellite—does not have any open faces. To satisfy requirements SYS-32 and SYS-33 (5.1), the monopoles will use the same deployable design used on CanX-2 (Figure 5.10(b)) such that no modiﬁcation would be needed on the XPOD-Triple deployment system. Instead of raising the deployables with motors or any other active electrical device, a torsion spring will be used to provide the required energy and actuation. The walls of the XPOD-Triple will constrain the deployables, eliminating the need for an actual release mechanism.

(a) Casing and cover (b) Deployment mechanism

Figure 5.10: Parts of the UHF radio

Power Board

The design of the power system was still uncertain at this stage in the design mainly because the GNB power board does not physically ﬁt within the 3U form factor. A few options that were considered included: Chapter 5. CanX-7 Structural Subsystem 52

• Re-designing the GNB power board

• Re-designing the CanX-2 power board

• Re-designing the NEMO power system

• Making use of the power system from the MESR (Mars Exploration Science Rover) project

It was desirable for the power board to have the same dimension as the HKC (167 × 20 mm) so that these components could be stacked together as in the GNB design, and then mounted onto the −Z panel, as in CanX-2. It was thought that for all the options, the re-designed system could be made to fit within a 167 × 20 mm board. Therefore, at PDR, the decision was to re-design the GNB power board and the modified power board was assumed to be the same shape and size as an HKC as this allocated area was deemed to be sufficient.

Magnetometer

Similar to the UHF antennas, the magnetometer used on a GNB is pre-deployed (Figure 5.11(a)).

As the XPOD-Triple does not have any open faces, the magnetometer used on CanX-7 must be deployable, and have a structural design similar to that of the CanX-2 magnetometer (Fig- ure 5.11(a)). The GNB magnetometer board is 22 × 42 mm whereas the CanX-2 magnetometer board is 20 × 30 mm. As a result, the magnetometer casing and its stowage area was enlarged to accommodate the larger dimensions. The magnetometer boom was also elongated by 40 mm to accommodate stowage over the S-band patch antenna (Figure 5.11(b)). The lengthened boom was not expected to cause vibration problems, given that it is stowed during launch.

(a) GNB and CanX-2 magnetometer (b) CanX-2 and CanX-7 magnetometer

Figure 5.11: CanX-7 magnetometer at PDR Chapter 5. CanX-7 Structural Subsystem 53

Magnetorquer

The wound magnetorquers used on GNB do not ﬁt within the 3U form factor. Printed torquers that are customizable in size were to be used for CanX-7 instead. They were to be placed on the

−X, −Y and −Z panels. The limiting factor on the size of the torquers was the space on the

−Z panel, as this panel of the three mentioned had the least amount of room on which to ﬁt the torquer (Figure 5.12). The largest torquer it could ﬁt is 76 × 61 × 6 mm and therefore CanX-7 was to use printed torquers of that size with holding bars to secure them in place (Figure 5.12).

Figure 5.12: Printed torquer design at PDR on Z− panel

5.2.3 Spacecraft Layout

The third aspect of the preliminary structural design is the layout of the spacecraft, whose internal and external conﬁguration at PDR is discussed in this section.

Internal

The main tray is the primary structure of CanX-7. It directly serves as a structural support to the S-band radio as well as to the imager and the COM DEV payload. Since the main tray has the S-band enclosure built-in, the S-band transmitter will be attached directly to the tray (Figure 5.13(a)). The width of the tray is 100 mm while that of the S-band enclosure is

86 mm. In order to leave room for wiring, the S-band enclosure was oﬀset by 2.5 mm from the Chapter 5. CanX-7 Structural Subsystem 54 center (Figure 5.13(b)). There was a 23.5 × 56 × 28 mm cutout on the tray where the imager was to be placed, and another cutout beneath the imager location for passage of imager wires

(Figure 5.13(a)). This was necessary because the UHF casing was placed above the S-band radio cover (which will mount to the tray above the S-band transmitter) covering access to the imager. Attachment of the UHF radio components (UHF casing, UHF Receiver and UHF cover) and the BCDR is shown in Figure 5.14(a).

(a) S-band Tx attachment (b) S-band enclosure oﬀset

Figure 5.13: S-band built into main tray

Aside from the main tray, the internal layout of the satellite includes the HKC, power board and magnetorquers. The HKC and power board will be stacked on top of each other and will mount onto the −Z panel (Figure 5.14(b)). The X, Y and Z magnetorquers will mount to the

−X, −Y and −Z panels, respectively (Figure 5.14(b)).

External

The components that are mounted on the outside of the solar panels of the satellite include:

• S-band patch antenna (×2)

• UHF monopole antenna (×4)

• UHF antenna guide (×4)

• Magnetometer Chapter 5. CanX-7 Structural Subsystem 55

(a) Tray attachment (b) Z− panel attachment

Figure 5.14: Internal layout at PDR

• Magnetometer boom guide

• Solar cells

These components were placed such that they would not interfere with the launch rails of the satellite. The S-band patch antennas were mounted on the +X and −X panels and were placed as close to the S-band transmitter as possible to minimize the coaxial cable length.

They were not placed at the same position on both +X and −X panels (Figure 5.15) because of other external components, such as the magnetometer and the solar cells. The positioning of the magnetometer, magnetometer boom guide, UHF antennas, UHF antenna guides and solar cells are shown in Figure 5.15. As a result of the magnetometer, only one solar coupon could be mounted on the +X panel while the rest of the long panels (−X, +Z and −Z) each have two solar coupons. The implication for power balance is that the duty cycles of certain components must be reduced. An external view of the satellite and a 180◦ rotation about the Y axis is shown in Figure 5.15. Chapter 5. CanX-7 Structural Subsystem 56

Figure 5.15: External view of the satellite (right) and 180◦ rotation (left)

5.3 Design Evolution

After the preliminary design review, the structural subsystem underwent another iteration of the design cycle. This was to address the issues and concerns that were raised during PDR.

Discussed in this section is the design evolution of certain components since the PDR that directly aﬀect the structural subsystem.

5.3.1 Drag Sails

The aspect of the drag sail design that concerns the structural subsystem is its attachment to the rest of the bus. At PDR, the four modules were to be stacked on top of each other in the conﬁguration shown in Figure 5.16(a) and placed on top of the bus, but a mature integration plan was not yet developed. After PDR, two additional plates were to be used in order to integrate the drag sails to the spacecraft (Figure 5.16(b)). Two drag sail modules were to mount to each of the mounting plates, which in turn attach to the four launch rails of the bus using four screws. Each plate was to be a maximum thickness of 10 mm and, as a result, an additional 20 mm was required for the drag sail stack. This inﬂuences the main tray design and the −Z panel as the feet of the launch rails require a further length increase of 20 mm. The current drag sail stack design and the design at PDR are both shown in Figure 5.16. Chapter 5. CanX-7 Structural Subsystem 57

(a) Drag sail module stack at PDR (b) Current drag sail module stack

Figure 5.16: Evolution of the drag sail module integration scheme

5.3.2 Secondary Payload

The secondary payload has undergone a major design evolution since the preliminary design review. The payload was changed from a Host and FemtoSat design to an Automatic Dependant

Surveillance-Based (ADS-B) Receiver. Both iterations of the payload are discussed presently.

Design I: Host and FemtoSat

At the time of PDR, uncertainty still existed regarding the secondary payload details. Its requirements were not well deﬁned and an interface control document (ICD) between the payload the spacecraft had not yet been created. Only the upper mass limit (1 kg) and the volume envelope (100 × 90 × 60 mm) of the payload were known. This payload maximum size and mass were very diﬃcult to accommodate in the 3U form factor.

A few weeks after PDR, a meeting with COM DEV was arranged to discuss the state of the secondary payload. During this meeting, it was learned that the secondary payload was to be provided by the Space Systems Dynamics and Control Lab of Ryerson University and was to consist of two components: an internal (Host) and an external part (FemtoSat), as shown in Figure 5.17(a)). This change greatly aﬀected the mounting location of other components on the satellite, the design of the primary structural parts and the spacecraft layout. Chapter 5. CanX-7 Structural Subsystem 58

The objective of the secondary payload is to test the capabilities of the FemtoSat (GPS, rate sensor, magnetometer and camera) on-orbit as well as to test the Bluetooth technology on board the Host. To begin the secondary payload operation, CanX-7 will operate in B-Dot mode with 0 torquer bias. A very high level secondary payload operation plan may be summarized in three steps, as follows.

1. The FemtoSat captures an image of the object of interest (Earth, moon, CanX-7) with

its camera.

2. The FemtoSat transmits the captured image wirelessly to the Host.

3. The image is transferred to the HKC and may be downloaded on subsequent satellite

passes.

(a) Secondary payloads (b) Mounting location

Figure 5.17: Post-PDR state of the secondary payloads and their mounting location

The Host is mounted internal to the satellite and has a volume envelope of 96 × 80 × 40 mm and a mass no greater than 350 g. As the initial allocated area for the secondary payload was greater than the size of the Host, the Host was able to be mount internally without any major design changes. It has ﬂanges which attach to the main tray and the −Z panel through eight screws. The top cover of Host rests on both ±X panels and is connected by four screws.

(Figure 5.17(b). Any communication between the secondary payload (including FemtoSat) and the bus will occur through the Host. The Host uses a 4-wire connection (data, clock, power Chapter 5. CanX-7 Structural Subsystem 59 and ground) to the rest of the satellite.

The FemtoSat was to be mounted external to the satellite on a boom and is meant to be a deployable. To accommodate this, the magnetometer was moved in-board so that the FemtoSat could use the magnetometer’s boom and deployable mechanism. The cutout on the panel where the magnetometer (28 × 35.8 × 6.5 mm) was originally recessed into had to be increased to accommodate the FemtoSat board that was to house a GPS, rate sensor, magnetometer and a camera. The largest size that could be allocated to the outer dimension of the FemtoSat enclosure was approximately 71 × 40.5 × 17 mm.

Originally, the mounting location of the FemtoSat was on the same panel as where the magnetometer had been (+X). At this location, the boom would be stowed atop the S-band patch antenna. In case the boom failed to deploy, the S-band antenna pattern would be aﬀected and a situation may have resulted where communication would be limited or even lost. To avoid boom placement on top of the S-band patch antenna, the FemtoSat deployable was moved to the adjacent panel (−Z). This was possible by reducing the dimension of the Host payload along the Z-axis by the depth of the FemtoSat enclosure. The mounting location of both the

Host and FemtoSat is shown in Figure 5.17(b).

As the deployable FemtoSat is bigger and heavier than the deployable CanX-2 magnetometer but is using the same CanX-2 deployable design, calculations were carried out to determine if the same spring was suﬃcient or if design changes were needed for correct sizing. An SFL engineer’s Master’s Thesis [12] was used as reference as it describes in detail the design and calculations that were performed on the CanX-2 deployable subsystem. The two metrics used to measure spring sizing for the FemtoSat were: 1) the time it took for the boom to fully deploy; and 2) the angular velocity of the FemtoSat at full deployment, which was 90◦. The deployable

N·m system uses a torsional spring with a torsional spring stiﬀness, KT , of 0.00375 rad as was the case for the analogous system on CanX-2. As this was a rotational system, a starting point was the rotational diﬀerential equation of motion:

X d2β(t) T = Iα ⇒ −K β(t) = I (5.1) T dt2

where KT is the torsional stiffness of the spring, I is the moment of inertia of the deployable Chapter 5. CanX-7 Structural Subsystem 60 about the axis of rotation and β(t) is the angular deflection of the torsion spring at time t. By solving the rotational differential equation for the system, the first metric could be calculated using the following equation:

r r KT I −1 β(t) β(t) = β0 cos t · ⇒ t = cos (5.2) I KT β0

where β0 is the initial position of the torsion spring (at time t = 0). Likewise for the second metric, the angular velocity (ω) of the FemtoSat in terms of time and angular position was also determined: r r K K ω(t) = − T β sin t T (5.3) I 0 I

s K ω(β) = T β2 − β(t)2 (5.4) I 0

The moment of inertia of the FemtoSat is an order of magnitude greater than the CanX-2 magnetometer; this is because the FemtoSat is approximately 150 g whereas the magnetometer is only 30 g. Also, the length of the FemtoSat boom is approximately 20 cm whereas the magnetometer boom is only 15 cm. Using these characteristic of the FemtoSat in the equations

rad above, it was calculated to deploy in about 1.3 seconds with an angular velocity of 2.13 s . As expected, this is much slower than the CanX-2 magnetometer deployment time of 0.46 seconds

rad with a faster angular velocity of 6.26 s [12]. Since two of the UHF deployable antennas were also on the same face as the FemtoSat, there was a concern that there could be some mechanical interference between the three deployables after ejection from the XPOD. As the UHF deployables employ the same design on CanX-7 as they did on CanX-2, their original CanX-2 spring characteristics were used in (5.2) to calculate a

UHF antenna deployment time of 0.15 seconds [12]. As this is an order of magnitude faster than it would take the FemtoSat to deploy, there is no real concern regarding the three deployables on the same face colliding with each other after ejection from the XPOD. Chapter 5. CanX-7 Structural Subsystem 61

Design II: Automatic Detection Surveillance-Based Receiver

As of writing this thesis, the secondary payload has changed from being a Host and FemtoSat to an Automatic Dependant Surveillance-Based (ADS-B) Receiver. As this is a recent change, the structural design of the ADS-B receiver was to be as close as possible to the design of the Host to minimize other structural changes within the rest of the satellite. To aid with the assembly and integration of the secondary payload with the rest of the structure, it was desirable that the secondary payload should be one monolithic piece. Also, the receiver was to mount externally to the CanX-7 structure. These factors resulted in the design of the secondary payload that may be seen in Figure 5.18(a).

(a) ADS-B payload (b) Mounted location

Figure 5.18: Current state of the secondary payload and its mounting location

5.3.3 Power System

In addition to the secondary payload, the design evolution of the power system also substantially inﬂuenced the structural subsystem. At PDR, the allocated power board was assumed to be the same size and shape as the HKC (Figure 5.19(a)). However, post-PDR, the decision was made to include a new modular power system designed for use in several upcoming missions.

This was desirable from a cost savings perspective, as the development cost could be shared between projects. The current power system consists of two µSPN (micro Switch Power Node) cards to deliver power to all the loads, one IFN (InterFace Node) and a POSIF (POwer System Chapter 5. CanX-7 Structural Subsystem 62

InterFace) board which controls and commands the power signals (Figure 5.19(b)).

(a) Assumed power board size at PDR (b) Current design

Figure 5.19: Evolution of the power system

5.3.4 Magnetometer

At PDR, CanX-7 had a deployable magnetometer which used a mechanism similar to that on

CanX-2. In its current form, the magnetometer has been moved in-board on the satellite and is mounted directly to the −Z panel. As previously explained in the secondary payload design evolution section ( 5.3.2), this was to accommodate the FemtoSat which needed to be deployed from the satellite.

5.3.5 Magnetorquers

Of the −X, −Y and −Z panels—the three panels to which the torquers are mounted—the

−Z panel had the least amount of room because of all the other components it needed to accommodate. In order to keep all the torquers the same size, they were sized according to this −Z panel room constraint. The thickness of the torquers needed to be a multiple of 2 mm as each layer of PCB on the printed torquers is 2 mm thick. The design changes of other components which mount to the −Z panel—such as the power system—resulted in decreasing the thickness of the torquers. As such, the thickness of the torquers decreased from 6 to 4 mm, post-PDR. Analysis presented by the attitude and orbit control subsystem indicated that the control authority provided by the 4 mm torquers were suﬃcient to meet attitude requirements. Chapter 5. CanX-7 Structural Subsystem 63

To attach the torquers to the panels, the alignment holes on the printed torquers (that were meant to align the diﬀerent layers of PCB) were used for mounting rather than the torquer holding bars, as shown in Figure 5.20.

(a) Torquer design at PDR (b) Current torquer design

Figure 5.20: Evolution of the magnetorquer design

5.3.6 Imager

After PDR, the imager was no longer included in the satellite as the post-PDR secondary payload—FemtoSat—-had a camera on-board and was intended to image the deployed sails for conﬁrmation. The increase in free volume allowed the radio stack to be shifted down and the battery to be mounted to the tray as shown in Figure 5.21. As a result of this change, the mass of the tray was reduced. However, due to the recent change in the secondary payload— currently an ADS-B receiver—an imager will be included in the satellite but not on the main tray. Speciﬁcs of this imager are still to be determined, but are beyond the scope of this thesis.

5.3.7 Primary Structural Components

The two primary structural components are the main tray and the −Z panel. The launch rails on both of these components were re-designed to be single solid pieces and therefore are no longer segmented. This eliminated the concern with segmented launch rails of their potential to become caught or stuck inside the XPOD during satellite ejection. Also, the feet of the launch Chapter 5. CanX-7 Structural Subsystem 64 rails were lengthened by another 20 mm (in addition to the 60 mm) from their PDR state in order to accommodate the two additional drag sail plates. The main tray was also modiﬁed for battery mounting (instead of being stacked on top of the radios). This was possible because of the reduction in the secondary payload volume, and the elimination of the imager. The design evolution of the main tray is shown in Figure 5.21.

Figure 5.21: Main tray design at PDR (left) and current design (right)

The −Z panel changed dramatically post-PDR mainly as a result of the power system design. Since the power system was to retain its shape, the mounting locations of the rest of the components required modiﬁcation. In addition to the power system, the −Z panel was to house the magnetometer, a magnetorquer, the HKC, and the separation switches. Mounting bosses of diﬀerent lengths were added so that parts attached to this panel could be mounted in a staggered fashion. The design evolution of the −Z panel is shown in Figure 5.22. Chapter 5. CanX-7 Structural Subsystem 65

Figure 5.22: −Z panel design at PDR (left) and current design (right)

5.4 Satellite Layout

As a result of the design changes that were made to the structural subsystem, the layout of the satellite changed since PDR. This was also a result of the suggestions from an SFL engineer,

Cordell Grant, who reviewed the solid model. The layout changed for more practical assembly of the satellite.

5.4.1 Layout of Main Tray

The main tray layout was changed such that the battery is now mounted directly to the panel due to the decrease in the initial secondary payload size. The main tray layout change since

PDR is shown in Figure 5.23.

Another major change to the main tray was the S-band and UHF radio layout. The S-band radio was rotated 180◦ so that the SMA connectors were at the −Y end of the satellite. The

S-band cover was discarded altogether and the bottom of the UHF enclosure was used in its Chapter 5. CanX-7 Structural Subsystem 66

Figure 5.23: Main tray layout at PDR (left), post-PDR (center), and current (right) place. An extension was built into the UHF enclosure such that it covered the SMA end of the S-band. The UHF enclosure was attached to the S-band enclosure using four screws: two screws were positioned next to the S-band SMA connectors while the other two were positioned at the back end of the UHF enclosure, also serving to secure the UHF cover. This second pair of screws pass through the UHF cover and enclosure and screw into the main tray behind the

S-band. The evolution of the radio layout post-PDR is shown in Figure 5.24. By implementing these changes, the new radio assembly is more eﬃcient in both volume and mass as some parts were eliminated and others were simpliﬁed. Further, by moving the S-band patch antenna to the −Y end of the satellite, the connection to the radio becomes practical for assembly by allowing passage of a person’s hand through the open end of the satellite. This was not possible previously, when the patch antenna was mounted toward the center of the satellite, meaning there was no feasible way in assembly to make the connection between the S-band radio and the antennas.

5.4.2 Layout of −Z Panel

The internal layout of the −Z panel is organized in diﬀerent layers as shown in Figure 5.25.

Attached closest to the panel itself are the separation switches, magnetometer, and the Z magnetorquer. The second layer is the power system, which is mounted to the panel through Chapter 5. CanX-7 Structural Subsystem 67

Figure 5.24: Radio assembly at PDR (left) and current layout (right) nine size zero screws. The size and positions of the screws on the boards were predetermined and thus were unable to be changed. As initially conceived, the HKC was to be mounted above the power system. However, the holes on the power system are not the same size as those on the HKC, making alignment impossible. Therefore, HKC speciﬁc mounting bosses were added.

All screws except the center screw on the +X side of the HKC are able to mount to the panel.

This exception is due to the power system located below, which would suﬀer from interference by this center screw. Therefore, an HKC attachment plate was created running the length of the HKC, which provides a center screw mount to which the HKC may be secured.

Figure 5.25: Layout of current −Z panel Chapter 5. CanX-7 Structural Subsystem 68

5.4.3 Layout of Other Internal Components

The only other internal components are the X and Y magnetorquers which are mounted to the

−X and −Y panels, respectively.

5.4.4 Layout of External Components

The external layout of the satellite consists of solar cell coupons, deployable UHF antennas, test port and S-band patch antennas which may be seen in Figure 5.26. Two of the long faces have two solar coupons each while the remaining four faces each have only one coupon. This allotment was driven by the surface area available on each of the panels. Both the main tray and the −Z panel have two deployable UHF antennas using similar mechanisms as on the CanX-2 satellite. In addition to the UHF antenna, the −Z panel also has the test port. The S-band patch antennas are on the ±X panels towards the −Y side of the satellite.

Figure 5.26: Current external layout

5.5 CanX-7 Mock-up

5.5.1 Prototype Structural Components

From the solid model of CanX-7, it may be seen that the spacecraft is highly space limited— despite being designed as volumetrically eﬃcient as possible. The main reason for this tightness of bus components within a 3U form factor is due to the secondary payload. When the primary Chapter 5. CanX-7 Structural Subsystem 69 payload (100 × 100 × 60 mm) is added, there is only approximately half of the entire satellite volume remaining for bus components. Further, these components are larger than those of their

CanX-2 counterparts, which contributes to an even more challenging volume arrangement. A main concern in the assembly of this satellite is the wiring harness design. The approach taken to reduce this risk was to manufacture a mock-up of the satellite to address issues not readily visible by analyzing the solid model itself. Structural components of the satellite were rapid-prototyped instead of machined for the following reasons:

• Mechanical drawings, which are time consuming to create, were not required for 3D

printing;

• Rapid prototyping was a order of magnitude less expensive than machining;

• Rapid prototyping was much faster than machining in terms of schedule; and

• Precision was not a crucial factor.

The parts were printed by U of T student machine shop in the Mechanical and Industrial

Engineering Department, as this source submitted the lowest quote for the order and only required STL ﬁles—a standard ﬁle format—for the components to be manufactured. The U of

T machining shop used a Dimension 3D Printer with a work envelope of 25.4 × 25.4 × 30.5 cm and build material of ABS plastic [19]. The solid model of two of the structural parts which exceeded the machine work envelope (i.e., the main tray and −Z panel) were split for 3D printing. They were each divided in a jagged shape so that more adhesive area would be available for their eventual re-joining (using glue) than a simple plane slice would oﬀer. An image of the primary structure solid model ﬁles and the manufactured components is shown in Figure 5.27. Other components that were also manufactured included the Host secondary payload, the S-band cover/UHF enclosure and the UHF radio cover.

After the parts were manufactured, some post-processing was required. The following steps were taken before the structural prototype was ready for assembly and integration.

First, the support material inside the holes of the printed parts had to be removed, as these parts were printed onto a breakaway support material instead of a soluble support material.

The support material is the layer onto which the machine deposits the build material to build up the required part (left of Figure 5.28). The right side of Figure 5.28 indicates how the −Z panel prototype was aﬀected when its breakaway support material was removed. Chapter 5. CanX-7 Structural Subsystem 70

(a) (b)

Figure 5.27: Solid model and prototype of structural parts for the CanX-7 mock up

(a) (b)

Figure 5.28: Parts printed on support material

The second main task was to strengthen both the holes—as they would be drilled and tapped—and the entire structure in general. Since the parts are made from ABS plastic, an

ABS cement solvent was applied to the material around the holes with a syringe. The needle was small enough to fit in the gaps between the walls of each hole and the plastic beads surrounding them, allowing the gaps to be filled. To strengthen an entire part, the structure was dipped in methylene chloride solution as suggested in reference [20]. The parts were then left to dry in the environmental room for several hours to stiffen and improve the surface finish. After drying, the split pieces of the main tray and the −Z panel were glued together using the ABS cement solvent. Chapter 5. CanX-7 Structural Subsystem 71

Third, the holes were drilled and tapped as the 3D printer does not print threaded holes.

As a result, the threaded holes in the solid model required readjustment to be simple holes with corresponding diameters to be tapped for Helicoil inserts or regular fasteners. Finally, after tapping, Helicoil threaded inserts were installed in most of the holes to reduce wear of the structure during assembly and disassembly. These inserts are mainly used at SFL to avoid soft thread damage by harder stainless steel screws. The “Helicoil Installation Guide” document [23] was used to properly install Helicoils into the manufactured parts.

5.5.2 Wiring Harness Design

After the post-processing, parts were ready for assembly and integration. As stated earlier, the main purpose of the mock-up was to determine wiring harness design. Additional reasons for building the satellite were to verify the design and determine if any other factors were not considered. The mock-up used non-ﬂight units for components that were already designed as well as blank printed circuit boards (PCBs) for components that were not yet ready. Spare hardware used for the mock-up included the S-band radio, UHF radio, HKC, magnetometer, two micro Switch Power Node (µSPN) cards and one InterFace Node (IFN) card. For the

POwer System InterFace (POSIF) board that encompasses the third part of the power system, a blank PCB was used.

To verify the design, a mechanical fit check of the components was performed by assembling the satellite (Figure 5.29(a)). During this assembly, most components attached to the spacecraft as expected. However, a few minor problems were encountered and subsequently fixed. For example, the middle stiffeners on the X panels interfered with the S-band radio enclosure and needed shifting by 1 mm to avoid this issue. Also, for several components, longer Helicoils and screws were used than originally designed. A partially assembled mock-up is shown in

Figure 5.29(b).

For the wiring harness, the CanX-7 wiring interconnect diagram was used [21]. Based on this diagram and the typical SFL colour code for wires [22], the appropriate selection was made for colour, gauge and quantity of wires. The wires were particularly configured to develop a mock-up as representative of actual flight configuration as possible, with the aim of determining the tie-wrap points, specific wire lengths and assembly procedure to be used for flight. As the Chapter 5. CanX-7 Structural Subsystem 72

(a) (b)

Figure 5.29: Mechanical ﬁt check and partial assembly of CanX-7 mock-up lengths of each wire were still undetermined, the wires used in the mock-up were prepared conservatively long as depicted in Figure 5.30.

Figure 5.30: Wiring harness for CanX-7 mock-up

5.5.3 Current Status

At the time of this writing, the mock-up has been utilized to verify the satellite design, determine and eliminate several previously unknown flaws, obtain a feeling for the amount of wiring needed Chapter 5. CanX-7 Structural Subsystem 73 inside the satellite and prepare an outline for the order of assembly. The wiring harnesses built are intended to be imported to the CanX-7 flat satellite (flatsat) which is to be built by the end of summer 2012. Before this stage, a few outstanding tasks remain to be completed with regards to the structural mock-up. These tasks include:

• Determining tie-wrap points on the satellite and length of wires;

• Designing a jig that supports the structure during integration and assembly;

• Utilizing the POSIF board to simplify the wiring and, in turn, develop its requirements;

and

• Creating a thorough assembly procedure for CanX-7.

5.6 Assembly and Disassembly

The assembly and disassembly of the CanX-7 satellite is a challenging task as the 3U form factor is very volume constrained. The order of assembly becomes extremely important as the scale of the satellite is on the same size as a human hand. Therefore, in designing the assembly procedure, considerations included the possibility of ﬁtting a hand inside certain spaces to insert fasteners or tighten screws. The plan was to create a thorough assembly procedure of CanX-7 during the building of the CanX-7 mock-up by iterating the wiring harness design several times.

In the interim, a top-level assembly procedure was developed and is provided below as well as in Figure 5.31.

1. Assemble main tray:

• Fasten BCDR;

• Fasten S-band radio;

• Assembly UHF radio;

• Mount UHF enclosure/S-band radio cover; and

• Fasten UHF radio cover.

2. Fasten the −Y solar panel to main tray.

3. Assemble −Z panel:

• Mount separation switches, magnetometer, and Z magnetorquer;

• Fasten power system; and Chapter 5. CanX-7 Structural Subsystem 74

• Mount HKC and HKC attachment plate.

4. Fasten −Z panel to −Y solar panel.

5. Integrate −X solar panel (might need temporary supports to hold the −Z panel and the

main tray in place).

6. Remove −Y panel:

• Connect S-band coaxial cables to patch antenna on the −X panel.

7. Integrate +X solar panel:

• Connect S-band coaxial cables to patch antenna on the +X panel.

8. Fasten −Y panel to −Z panel and main tray.

9. Integrate ADS-B receiver.

10. Integrate drag sail modules.

11. Integrate +Y panel.

Figure 5.31: Spacecraft-level assembly procedure for CanX-7 Chapter 5. CanX-7 Structural Subsystem 75

5.7 Mass Budget

The mass budget has been updated consistently throughout the CanX-7 design evolution. The most up-to-date mass budget for CanX-7 is summarized in Table 5.1. This table compares the mass of the satellite at two points in time: during the early stages of the mission—when CanX-7 was a conceptual design—and at its current state. A detailed mass budget of the current design is provided in Table 5.2. During the initial design, the target mass was based on the ﬁnal mass of CanX-2 of 3.5 kg as CanX-7 will use the same ejection system—an XPOD-Triple. Currently, a target mass of 4 kg is used as the XPOD-Triple will be qualiﬁed for 4 kg. At present, there is a 16% margin employed—which has been increased since the initial mass estimate. This margin is expected to increase as the design matures and requires less contingency. The trend for the majority of subsystems is a decrease in mass. Outlined below are the major changes that were made to certain subsystems during the design process which helped reduce the overall satellite mass:

• Structural (Mass decrease):

– The main tray is lighter than the initial design;

– The +Z panel was eliminated (it is built into the main tray); and

– Panels and main tray are to be machined from magnesium instead of aluminum.

• AOCS (Mass decrease):

– Sun sensors were no longer required; and

– ADCC (Attitude Determination and Control Computer) was excluded.

• Power (Mass decrease):

– A total of nine solar cells are to be used instead of 10; and

– CanX-2 wiring harness mass is used for estimates (instead of GNB) as CanX-7 has

a 3U form factor.

• Communications (Mass increase):

– UHF enclosure is required (as it is unable to be built into the main tray).

• Payloads (Mass decrease):

– COM DEV payload mass is approximately 500 g (instead of 1 kg); and

– Contingency on drag sail modules has decreased (sail design has matured). Chapter 5. CanX-7 Structural Subsystem 76

Table 5.1: Mass budget comparison of conceptual and current designs Conceptual Design Current Design Subsystem Mass (g) Fraction (%) Mass (g) Fraction (%) Trend Structural 1,257 28 852 25 Decrease Thermal Control 45 1 49 1 Increase AOCS 186 4 93 3 Decrease Power 760 17 636 19 Decrease Computer 63 1 60 2 Decrease Communications 349 8 405 12 Increase Payloads 1,863 41 1,219 36 Decrease Integration 45 1 33 1 Decrease Total 4,568 100 3,347 100 Target 3,500 – 4,000 – Margin -1,068 -31 653 16

5.8 Structural Analysis

The purpose of structural analysis is to determine the structural integrity of the design. A structural analysis of the CanX-7 satellite was carried out to ensure that the satellite would withstand the loads and vibration experienced during launch as expressed by the launch vehicle

(LV) provider. A Finite Element Model (FEM) of the CanX-7 satellite was created and the launch environment was simulated using NX software. The FEM was used to determine the natural modes of the satellite and maximum stresses and deﬂections of the structure during worst-case loading conditions. This section describes how the FEM was created, the boundary conditions used to simulate the satellite within the XPOD during launch and the results from the modal and quasi-static analysis.

5.8.1 Background

A finite element model (FEM) is a representation of a real object composed of a network of nodes with finite elements between them that describe the geometry and stiffness characteristics of the entire part. A finite element analysis (FEA) can be performed on the FEM using various types of solvers that are currently on the market. For CanX-7, the NX Nastran Solver was used.

In creating an FEM, it is desirable to minimize the number of elements in the model to Chapter 5. CanX-7 Structural Subsystem 77

Table 5.2: Detailed mass budget of the current design

Individual Unit CanX-7 Mass Estimated Contingency Total Contingency Total Frac- Mass Mass Mass Mass Mass tion Component Status (g) (g) (g) No. (g) (g) (%) Structural Subsystem 6.0 851.8 25 Main Tray and Rails (with Z+ Panel) M1 241.7 12.1 253.8 1 0.0 253.8 X− Solar Panel M1 94.0 4.7 98.7 1 0.0 98.7 X+ Solar Panel M1 94.0 4.7 98.7 1 0.0 98.7 Y − Solar Panel M1 34.0 1.7 35.7 1 0.0 35.7 Y + Solar Panel M1 34.0 1.7 35.7 1 0.0 35.7 Z− Solar Panel (includes Rails) M1 188.0 9.4 197.4 1 0.0 197.4 Screws M1 0.5 0.0 0.5 224 5.4 118.7 Spacers M1 0.2 0.0 0.2 54 0.6 13.1 Thermal Control Subsystem 8.0 48.7 1 S-band Thermal Strap M1 3.2 0.2 3.4 1 0.2 3.4 Battery Gaskets M1 7.4 0.4 7.8 1 0.4 7.8 Thermal Tapes E 5.0 1.25 6.3 6 7.5 37.5 Attitude Control Subsystem 4.2 93.2 3 Magnetorquer M1 28.0 1.4 29.4 3 4.2 88.2 Magnetometer (in-board) M2 5.0 0.0 5.0 1 0.0 5.0 Power Subsystem 24.3 635.9 19 Batteries M2 105.0 0.0 105.0 1 0.0 105.0 Battery Collars M2 13.2 0.0 13.2 2 0.0 26.4 BCDR M2 37.3 0.0 37.3 1 0.0 37.3 Power Board E 75.0 18.8 93.8 1 18.8 93.8 Solar Panel (with Wires) M2 28.5 0.0 28.5 9 0.0 256.5 Wiring Harness (with Coaxial Cables) M1 111.4 5.6 117.0 1 5.6 117.0 Computer Subsystem 0.0 60.0 2 HKC M2 60.0 0.0 60.0 1 0.0 60.0 HKC Attachment Plate M1 18.5 0.9 19.4 1 0.9 19.4 Communications Subsystem 6.3 404.9 12 S-band Transmitter (Core) Board M2 55.0 0.0 55.0 1 0.0 55.0 S-band LNA (Output) Board M2 57.0 0.0 57.0 1 0.0 57.0 S-band Output Board Johnson Con- nectors M2 8.0 0.0 8.0 2 0.0 16.0 S-band Shielding Plate M2 24.7 0.0 24.7 1 0.0 24.7 S-band Patch Antennas M2 11.7 0.0 11.7 2 0.0 23.4 UHF Rx M2 60.0 0.0 60.0 1 0.0 60.0 UHF Rx Johnson SMA Connectors M2 5.0 0.0 5.0 4 0.0 20.0 UHF Antennas (Deployables) M2 4.0 0.0 4.0 4 0.0 16.0 UHF Rx Enclosure M1 74.4 3.7 78.1 1 3.7 78.1 UHF Radio Cover M1 52.1 2.6 54.7 1 2.6 54.7 Payloads 36.6 1,219.4 36 Drag Sails M1 150.0 7.5 157.5 4 30.0 630.0 Drag Sail Plates M1 50.0 2.5 52.5 2 5.0 105.0 Secondary Payload (Host) M2 350.0 0.0 350.0 1 0.0 350.0 Secondary Payload (FemtoSat) M2 100.0 0.0 100.0 1 0.0 100.0 Secondary Payload (FemtoSat De- ployable Mechanism) M1 32.8 1.6 34.4 1 1.6 34.4 Integration 1.0 33.0 1 Total 86.0 3,347.0 100 Target – 4,000.0 – Margin – 653.0 16.3 Chapter 5. CanX-7 Structural Subsystem 78 obtain a quick solution. For complex geometries, the part may be split into diﬀerent regions with corresponding meshes to accurately represent the structure while keeping the number of elements low. For example, more elements should be used in the regions where stress concentration is expected, and fewer elements for areas where the stress is not expected to vary greatly.

There are diﬀerent types of elements that may be used to create the FEM. The choice of element type for a part should be one that is able to closely replicate the object. There are lumped mass (0D mesh), shell (2D mesh), and solid (3D mesh) elements. The 0D elements are mainly used for objects whose mass cannot be neglected but whose stiﬀness may be neglected, while the 2D elements are used to model thin sheets. There are three types of solid elements: bricks, regular order tetrahedra, and higher order tetrahedra, which are each made up of four,

ﬁve and six nodes, respectively. The two latter element types are used for objects that do not have a simple shape.

5.8.2 Finite Element Modeling

The CanX-7 structural FEM was created in NX by ﬁrst importing the solid model from Solid-

Edge software that was primarily used to design the spacecraft. The entire spacecraft FEM was constructed in sections—as the parts had different geometries—and different meshes were used accordingly. The primary structural parts (−Z, ±X, ±Y panels and main tray) were de-featured (removing mounting holes and fillets) to simplify the parts as much as possible while maintaining their overall shapes to provide accurate results. Removed holes were not from locations on the satellite providing structural support and therefore large stress concentrations in those areas were not expected. Further, by removing fillets, the results from the model would be more conservative as the modeled stress concentrations in these areas would be greater than reality. The resulting idealized parts were linked to their original counterparts so that any change to the original part would be reflected in its corresponding idealized part, when updated. The idealized parts—which were to be meshed with regular order elements— were split into multiple regions. This way, mesh control could ensure the most appropriate mesh was used on each different region. For example, the meshes around the main satellite fastener holes were made to be finer than the mesh of the rest of the structure, as large stress Chapter 5. CanX-7 Structural Subsystem 79 concentration is expected around the fastener locations.

The FEM of CanX-7 consisted of solid (3D Tetrahedral mesh) elements, shell (2D mesh) elements and lumped mass (0D mesh) elements. The main tray and the −Z panel were created using 3D elements as they were geometrically sophisticated and therefore could not be simpliﬁed to be 2D shell elements. The screws from the main tray and −Z panel are received by the ±X panels. To model this connection which holds the satellite together, the cross braces of the

±X panels—the receiving end of the satellite fastener—were also created out of 3D elements, while the rest of the ±X panel was created out of 2D shell elements. Likewise, the drag sail attachment plates—which attach to the rest of the satellite—were also modeled using 3D elements. Similar to the main tray, −Z, and ±X panels, the ±Y panels provide stiﬀness to the satellite. However, they were created out of 2D shell elements as they are relatively simple shapes and are not at the receiving end of any screws binding the satellite together.

For the rest of the components, such as the battery, magnetometer and radios, lumped mass elements were used as the stiﬀnesses of the components could be neglected but their masses required consideration. For each of these parts, the center of mass determined from SolidEdge was used to place the 0D element at the correct position on the model. These components are connected to the rest of the structure through fasteners. To attach the lumped masses to their mounting points, a 1D node-to-node connection with rigid body element (RBE3) was used, as the RBE3 was speciﬁcally created for point mass distribution. The masses assigned to these components included not only the component masses, but also the masses of the fasteners that attach the components to the structure. Lumped mass elements were also used for components attached directly to the surface of the satellite, such as solar cells and the S-band patch antennas.

Even though they are not fastened at speciﬁc points, for simplicity, such lumped mass elements were modeled as attached at only a few points evenly distributed across their contact surfaces.

This is more realistic than evenly distributing attachment across the entire structure as the solar cells and the S-band patch antennas do not span the entire structure. An image of the

CanX-7 FEM is shown in Figure 5.32(a). Chapter 5. CanX-7 Structural Subsystem 80

(a) FEM of the CanX-7 satellite (b) BC imposed on the model

Figure 5.32: Finite element model and the imposed boundary conditions of CanX-7

5.8.3 Boundary Conditions

During launch, CanX-7 will be constrained within the XPOD by the feet of the four launch rails.

These rails are prevented from translating in the X and Z directions, which is a constraint on all eight feet. Since the satellite will compress in the Y direction under the acceleration during launch, only one end of each launch rail is constrained in the Y direction.

Contact surfaces within the satellite were also modeled for realistic behaviour. Panels rest along the edges of the structure as opposed to being constrained only at fasteners to which they are connected. As such, panels will not overlap in any regions with their contact surfaces. The edges of the main tray and the −Z panel are in contact with the ±X panels, the −Y panel and the two drag sail plates. The +Y panel is not attached to the main structure but is only in contact with its closest drag sail plate.

In addition to the boundary conditions imposed on the model (Figure 5.32(b)) to represent its conﬁguration during launch within the XPOD, the CanX-7 FEM was also subject to certain loading conditions to simulate the launch environment. Since the launch is not ﬁnalized, the structure has to withstand the worst-case loading conditions of two potential launch vehicles

(LV) that are being considered for CanX-7: Dnepr and PSLV (Polar Satellite LV). The following

Table 5.3 includes certain structural requirements of both LVs [26] for spacecraft acceptance level tests [16], [17]. Chapter 5. CanX-7 Structural Subsystem 81

Table 5.3: Launch vehicle providers’ acceptance level test requirements Parameters Dnepr PSLV Frequency (Hz) 20 90 8.2 g compression 7 g compression Longitudinal (axial) load 1.0 g tension 2.5 g tension 2.5 g compression 0.8 g compression Lateral load 2.5 g tension 1.7 g tension Ultimate load factor – 1.25 Test acceleration 8.2 g 7 g × 1.25 = 8.75 g

The worst-case conditions of either launch vehicle scenario were used to perform the following two analyses on CanX-7. The first analysis was predicting the first natural frequency of the satellite through modal analysis. The second was a quasi-static analysis to determine maximum stresses and deflections within the satellite.

5.8.4 Modal Analysis

For modal analysis, the satellite was subjected to the boundary condition of being within the

XPOD. This analysis was carried out to ensure that the ﬁrst natural frequency of the satellite is above the 90 Hz as required by PSLV, which is the worst-case between the two launch vehicles, to avoid dynamic coupling between the spacecraft and the LV. The ﬁrst unconstrained natural frequency of CanX-7 was 337 Hz, which is well above the maximum frequency of 90 Hz. This mode is the lateral displacement of the secondary payload, which is a concentrated mass of

500 g. (Figure 5.33).

5.8.5 Quasi-Static Analysis

The second type of analysis that was performed was to determine the maximum stresses and deﬂections of the satellite. The worst-case quasi-static launch load of 8.75 g was used as the PSLV test requirement envelops the Dnepr test requirement. Since the orientation of the satellite within the LV is unknown, it was subjected to the acceleration of 8.75 g in all three axes simultaneously. The highest stress observed was around 80 MPa which is below the yield stress of magnesium (130 MPa), and the maximum deﬂection observed was less than 0.03 mm Chapter 5. CanX-7 Structural Subsystem 82

Figure 5.33: First natural frequency of CanX-7 on the −X solar panel. Both of these results were considered acceptable. Figure 5.34 shows the deformed model when it was subjected to the PSLV worst-case acceleration in all three axis simultaneously.

Figure 5.34: Satellite stress spectrum under the worst-case acceleration

5.8.6 Hand Calculations

Hand calculations were carried out to ensure that fasteners within the satellite are able to withstand the launch loads. This was done because fasteners are complicated to model in the

FEM. Two calculations were performed:

1. Shear stress in the sixteen main fasteners that hold the entire satellite structure together Chapter 5. CanX-7 Structural Subsystem 83

2. Shear stress in the smallest fasteners within the satellite—ones that hold the power system

in place

The fasteners used on CanX-7 are type 316 stainless steel. Based on the “Stainless Steel

Fastener” designer handbook [25], the allowable shear stress for fasteners with no threads in the shear plane is 60% of the minimum tensile strength divided by a safety factor of 3. The minimum tensile strength of stainless steel is 515 MPa. This corresponds to an allowable shear stress of 103 MPa. For fasteners that have threads in the shear plane, as is is the case for the ones on CanX-7, 70% of the allowable shear stress is used. Therefore, the allowable shear stress for the fasteners on CanX-7 is 72.1 MPa. The two following calculations are to ensure that the shear stress seen by the fasteners on CanX-7 are below their allowable shear stress of 72.1 MPa.

Shear Stress in Satellite Fasteners

The entire satellite structure is held together by sixteen 12.7 mm #4-40 type 316 stainless steel fasteners. The shear stress, τ, experienced by each of the 16 fasteners was calculated from the shear stress equation, V τ = (5.5) A where V is the shear force experienced by the object during launch and A is the surface area of the fasteners holding the object in place. The surface area of a fastener (in British Imperial units) is calculated using the following formula from the “Stainless Steel Fastener” designer handbook [25]: 0.97432 A = 0.7854 D − (5.6) s n where D is the major diameter of the fastener in inches and n is the number of threads per inch.

For simplicity, the calculations were converted to metric. For a single #4-40 screw, which

−6 2 corresponds to D = 2.84 mm and n = 40 threads per inch, the surface area, As = 3.89 × 10 m . The force experienced by the CanX-7 satellite (3.35 kg)under loading condition of 8.75 g, is Chapter 5. CanX-7 Structural Subsystem 84

287.6 N. The shear stress experienced by each of the 16 fasteners is

V τ = A 287.6 = 16 × 3.89 × 10−6 = 4.6 MPa

Therefore, the screws are suﬃcient to withstand the launch environment as their expected shear stress of 4.6 MPa is much less than their allowable shear stress of 72.1 MPa.

Shear Stress in Power System Fasteners

Another hand calculation was performed on the nine #0-80 type 316 stainless steel screws that hold the power system in place, as these are the smallest-sized screws on the satellite. The force experienced by the power system, which is approximately 100 g in mass, under loading condition of 8.75 g, is 8.6 N. The surface area of a #0-80 screw calculated using the formula in

−6 2 (5.6) with D = 1.52 mm and n = 80, is As = 1.16 × 10 m . The shear stress experienced by the power system is

V τ = A 8.6 = 1.16 × 10−6 = 7.4 MPa

This is less than the allowable shear stress of 72.1 MPa for a type 316 stainless steel screw.

Further, since the power system is mounted with nine #0-80 screws, the shear stress experienced by each is ideally one ninth of the total 7.4 MPa. Therefore, like the main satellite fasteners, the smallest screws on the satellite are also suﬃcient to withstand the launch environment.

Even though screws are not meant to withstand shear loads, the previous analysis indicates that the fasteners could handle the expected shear stresses. This is part of the bottom-up design practice: the satellite is ﬁrst designed with what is practically possible with limited mass and volume and the structural analysis is then carried out to validate the design. Chapter 5. CanX-7 Structural Subsystem 85

5.9 Future Work

The CanX-7 structural subsystem has evolved signiﬁcantly over the course of this Master’s thesis. At the time of writing, the structural design is acceptable as indicated by analysis, and a high-level assembly procedure has been developed. A Critical Design Review (CDR) of

CanX-7 is planned to be held in early Fall 2012. As such, the following are outstanding tasks to be accomplished in the near future for the structural subsystem but are beyond the scope of this thesis:

• Develop a mature integration strategy for the drag sail modules;

• Create a detailed satellite assembly procedure; and

• Document and present the structural subsystem design at the critical design review.

After CDR, the plan is to accomplish the tasks listed below for the structural subsystem to be prepared for launch. These are for future work and not part of the author’s thesis.

• Send the primary structural components out for manufacturing for ﬂight;

• Lay down solar cells and apply thermal tape;

• Perform “dirty” satellite integration (i.e., not ready for launch); and

• Assemble the satellite for ﬂight. Chapter 6

Conclusion

This thesis described preliminary system development, deployment detection subsystem design and detailed structural design and analysis for the CanX-7 spacecraft. Over the course of two years, this mission has evolved signiﬁcantly from its conceptual design to near-readiness for

Critical Design Review (CDR). During this period and as part of the CanX-7 team from its beginning, the author has gained invaluable experience and understanding of the microspace approach to designing nanosatellites.

Performing initial system development work helped evaluate various bus designs for CanX-7 at a time when its funding situation was in a state of ﬂux. Carrying forward multiple (in this case, three) designs in parallel is a challenging but necessary task when a mission is in its early stages. The author learned the importance of cost as a driving factor in the decision-making process of real missions. As a result, cost considerations essentially dictated that CanX-7 would have a triple CubeSat (3U) form factor.

Due to the nature of design, the spacecraft underwent multiple iterations with some components or systems either being added or removed. An example of this was the Deployment

Detection Subsystem (DDS) worked on by the author. The DDS evolved from using a Com- merical Oﬀ-The-Shelf (COTS) camera, to being eliminated due to redundancy of function with the FemtoSat secondary payload (whose details were initially unknown) to currently being re- instated. Other components—such as sun sensors, rate sensors and reaction wheels—initially included in the CanX-7 design were later eliminated as it became clear that such items were unnecessary to meet mission requirements. Changes such as these are inherent to the design

86 Chapter 6. Conclusion 87 process, and these specific examples taught the author the importance of flexibility in mission definition prior to configuration freeze.

The author has also learned a great deal about the process of subsystem design through work on the structural layout of CanX-7. Although the preliminary structural subsystem design met all applicable requirements, a better design was achieved after insight was gained during preliminary design review (PDR). Feedback from both PDR and consultation with an experienced mechanical engineer was implemented to improve design quality. To verify the improved design, a structural mock-up of the CanX-7 satellite was built. This served to determine the wiring harness design, perform a mechanical-ﬁt check and develop an assembly procedure. A

ﬁnite element analysis (FEA) was then carried out to ensure structural integrity of the CanX-7 spacecraft design, which completed one iteration of the design cycle for the structural subsystem.

Over the past two years, the author met all her objectives stated in Section 1.4. The work summarized in this Master’s thesis plays an important role in the success of the CanX-7 mission, which will contribute to solving the global problem of space debris. By demonstrating its compact and bus-independent de-orbiting technology for nano-class satellites, CanX-7 will secure a safer frontier for future generations of scientists and engineers embracing the microspace philosophy. Bibliography

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