Canadian Advanced Nanospace eXperiment 7 (CanX-7) Mission ... · Abstract...

Canadian Advanced Nanospace eXperiment 7 (CanX-7) Mission Analysis,

Payload Design and Testing

by

Barbara Shmuel

A thesis submitted in conformity with the requirementsfor the degree of Master of Applied Science

Graduate Department of UTIASUniversity of Toronto

Copyright c© 2012 by Barbara Shmuel

Abstract

Canadian Advanced Nanospace eXperiment 7 (CanX-7) Mission Analysis, Payload Design and Testing

Barbara Shmuel

Master of Applied Science

Graduate Department of UTIAS

University of Toronto

2012

A deorbiting drag device is being designed and built by the University of Toronto Institute for Aerospace

Studies/Space Flight Laboratory (UTIAS/SFL) to be demonstrated on the Canadian Advanced Nanospace

eXperiment 7 (CanX-7) satellite. CanX-7 will address the growing issue of space debris by designing a

drag sail device that will be demonstrated for cubesat-sized satellites. Mission analysis done to ensure the

drag device functions properly and deorbits within the required lifetime is performed while varying different

properties such as drag coefficient, effective drag area, and solar cycle variations. The design evolution of

the device is documented and the chosen design, along with several stages of prototyping, is described. The

individual components that make up the device are described as are preliminary numerical analyzes. Finally,

the test plan required for the device is described with several deployment experiments and risk reduction

tests documented.

ii

Acknowledgements

I would first like to thank Dr. Robert E. Zee for giving me the opportunity to work on the CanX-7 project,

both as a summer and Masters student. This experience has been amazing and I am very grateful for the

chance to be a part of it.

I would also like to thank all the staff and students at SFL for providing their wisdom and knowledge.

All your out-of-the-box ideas for drag sails really helped and made the difficult design process that much

more enjoyable.

Finally to my family and friends, I would not have been able to get through the last two years without

your love and support. Thank you for always being there and helping me along my journey.

iii

Contents

1 Introduction 1

1.1 Space Flight Laboratory . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.2 Space Debris Accumulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

1.3 Canadian Advanced Nanospace eXperiment 7 . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

1.4 Scope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

2 Background: Space Debris Cause and Mitigation 7

2.1 The Space Debris Problem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

2.2 Significance of Deorbiting Satellites . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.3 Mitigation Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.3.1 Active Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.3.2 Passive Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

2.4 Deorbiting Device for CanX-7 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

3 Mission Analysis 13

3.1 Lifetime Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

3.1.1 Drag Sail Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

3.1.2 Interpreting the Lifetime Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

3.1.3 Drag Coefficient . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

3.1.4 Effective Area . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

3.1.5 Solar Effects on Lifetime . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

3.1.6 Time-Area Product . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

3.1.7 Error Approximations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

3.1.8 Expected Lifetimes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

3.2 Drag Sail On-Orbit Operations Plan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

iv

3.2.1 Orbit Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

3.2.2 Drag Sail Deployment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

3.2.3 On-Orbit Experiments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

4 Drag Sail Device 36

4.1 Initial Drag Sail Concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36

4.1.1 Deorbit Capabilities . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37

4.1.2 Dimensions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38

4.2 Design Evolution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

4.2.1 Folded Frame . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40

4.2.2 Coiled Frame . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40

4.2.3 Double Coiled Frame . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41

4.2.4 Initial Drag Sail Folding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42

4.3 Updated Drag Sail Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45

4.3.1 Design Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46

4.3.2 Housing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48

4.3.3 Sail Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53

4.3.4 Booms . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58

4.4 Prototyping . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59

4.5 Sail to Boom Attachment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63

4.6 Change of Device Dimensions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64

4.7 Drag Sail Placement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65

4.7.1 Stacked Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65

4.7.2 Tiled Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66

4.8 Structural Capabilities . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67

4.8.1 Drag Force . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67

4.8.2 Deployment Kinematics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68

4.9 Thermal Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

4.9.1 Relaxation Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

4.9.2 Deployed Drag Sail . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

5 Drag Sail Testing 73

5.1 Qualification Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74

5.1.1 Standard Deployment Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74

v

5.1.2 ReWind Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75

5.1.3 Repeated Deployment Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75

5.1.4 Vibration Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76

5.1.5 Thermal Vacuum Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76

5.1.6 Radiation Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77

5.2 Acceptance Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77

5.2.1 Standard Deployment Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77

5.2.2 Rewind Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78

5.2.3 Thermal Ambient Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78

5.2.4 Thermal Shock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78

5.3 Long-Term Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79

5.3.1 Long Term Storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79

5.3.2 Long Term Vacuum Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79

5.4 Drag Sail Deployment Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80

5.5 Risk Reduction Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81

5.5.1 Boom Exclusion Zone . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81

5.5.2 Tear Propagation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 84

6 Conclusion 90

vi

List of Tables

3.1 Small Error in the Effective Area Lifetime Calculations Example . . . . . . . . . . . . . . . . 21

3.2 Expected Lifetimes for a 2013 Launch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

4.1 Summary of Calculated Properties of Different Folds . . . . . . . . . . . . . . . . . . . . . . . 44

4.2 Sail Material Trade Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 54

4.3 Boom Material Trade Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58

vii

List of Figures

1.1 CanX-2 Satellite (Courtesy SFL) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

1.2 GNB and NEMO Satellites (Courtesy SFL) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

1.3 CanX-7 Satellite with Deployed Drag Sails . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

2.1 Space Debris Visualization [14] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

2.2 Solar Sail [15] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

2.3 GOLD Rigidizable Device [16] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

2.4 Electrodynamic Tether [17] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

3.1 Required Drag Area at Different Altitudes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

3.2 Comparison of Different Drag Coefficients . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

3.3 3U Effective Area Lifetime . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

3.4 GNB Effective Area Lifetime . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

3.5 NEMO Effective Area Lifetime . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

3.6 Lifetime for a 3U form factor versus Launch Year . . . . . . . . . . . . . . . . . . . . . . . . . 21

3.7 Lifetime for a GNB form factor versus Launch Year . . . . . . . . . . . . . . . . . . . . . . . 22

3.8 Lifetime for a NEMO form factor versus Launch Year . . . . . . . . . . . . . . . . . . . . . . 22

3.9 Lifetime for a NEMO form factor versus Launch Year, Updated Solar File . . . . . . . . . . . 23

3.10 Solar Flux Predictions over Several Solar Cycles [28] . . . . . . . . . . . . . . . . . . . . . . . 24

3.11 Orbit Altitude versus De-orbiting Time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

3.12 Time Area Product for a NEMO Satellite with and without a Drag Device . . . . . . . . . . 26

3.13 Variations amongst STK Solar Flux Files . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

3.14 Comparison of Different Ballistic Coefficients . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

3.15 Comparison of Different Atmospheric Models . . . . . . . . . . . . . . . . . . . . . . . . . . . 30

3.16 Expected CanX-7 Lifetime for a 2013 Launch for Different Drag Areas . . . . . . . . . . . . . 32

viii

4.1 Initial Drag Sail Concept (Courtesy of SFL) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37

4.2 Maximum Drag Sail Cross Sectional Area . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

4.3 Folded Frame Concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40

4.4 Coiled Frame Concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41

4.5 Double Coiled Frame Concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41

4.6 Butterfly Fold [32] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

4.7 Leaf Fold [32] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

4.8 Circular Fold [33] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

4.9 Miura-Ori [34], [35] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

4.10 Original Drag Sail Concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45

4.11 Drag Sail Dimensions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46

4.12 Drag Sail Device Placement on a 3U Satellite . . . . . . . . . . . . . . . . . . . . . . . . . . . 47

4.13 Drag Sail Device Placement on a GNB Satellite . . . . . . . . . . . . . . . . . . . . . . . . . . 47

4.14 Initial Drag Sail Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48

4.15 Old Drag Sail Concept with Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 49

4.16 Door Tray . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50

4.17 Close-up of Door Hinge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51

4.18 Drag Sail Housing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52

4.19 Drag Device Reel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52

4.20 Drag Sail Device Housing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53

4.21 Problems with Miura-Ori Deployment Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

4.22 Drag Sail Material Fold . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56

4.23 First Stage Folding Machine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56

4.24 Resulting Fold after First Stage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57

4.25 Second Stage of Folding Machine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57

4.26 COTS Boom . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59

4.27 Round Small Prototype . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60

4.28 Round Large Prototype . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60

4.29 Large Reel Prototype . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61

4.30 Large Reel Prototype in a Stowed Configuration . . . . . . . . . . . . . . . . . . . . . . . . . 62

4.31 Coiled Boom Concept Validation - Deployment Sequence . . . . . . . . . . . . . . . . . . . . 63

4.32 Sail to Boom Initial Attachment Concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64

4.33 New Drag Sail Housing (Courtesy of SFL) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65

ix

4.34 Drag Sail Placement on a 3U (Courtesy of SFL) . . . . . . . . . . . . . . . . . . . . . . . . . 66

4.35 Drag Sail Placement on a GNB (Courtesy of SFL) . . . . . . . . . . . . . . . . . . . . . . . . 66

4.36 Drag Sail Placement on a NEMO (Courtesy of SFL) . . . . . . . . . . . . . . . . . . . . . . . 67

4.37 Drag Force Experienced by the Drag Sail at Different Altitudes . . . . . . . . . . . . . . . . . 68

4.38 Relaxation Curve for Carbon Steel [48] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

4.39 Hubble Space Telescope Booms Bending [49] . . . . . . . . . . . . . . . . . . . . . . . . . . . 72

5.1 Re-foldability Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80

5.2 Boom Excursion Deployment Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82

5.3 Example of Incomplete Sail Material Deployment . . . . . . . . . . . . . . . . . . . . . . . . . 82

5.4 Boom Excursion Deployment Test, No Sail Material . . . . . . . . . . . . . . . . . . . . . . . 83

5.5 Stress versus Fatigue Life for Kapton [52] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 84

5.6 Kapton Endurance Limit [52] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 85

5.7 Tear Propagation: Centre and Edge Tear Oscillations . . . . . . . . . . . . . . . . . . . . . . 86

5.8 Tear Propagation: Edge Tear Oscillations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 86

5.9 Tear Propagation Test Coupon . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 87

5.10 Tear Propagation Test Setup . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 87

5.11 Tear Propagation Test Setup - Connection to Motor . . . . . . . . . . . . . . . . . . . . . . . 88

5.12 Hall Effect Sensor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 89

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Chapter 1

Introduction

The University of Toronto Institute for Aerospace Studies’ Space Flight Laboratory (UTIAS/SFL) is at the

forefront of new, innovative small satellite technology. Building small satellites on an aggressive schedule

at low cost is the microspace philosophy under which the lab operates. The UTIAS Space Flight Lab is

developing leading-edge satellites for research and development in space and providing platforms for mission

specific goals. The Canadian Advanced Nanospace eXperiment 7, CanX-7, will be a new mission for SFL

and focuses on the growing space debris problem that is quickly becoming a concern for small satellites. This

thesis will introduce the work done on the CanX-7 primary payload and supporting analyzes.

1.1 Space Flight Laboratory

The UTIAS Space Flight Laboratory was founded in 1998 and has grown into a lab that is developing and

building some of the most advanced small satellites in Canada and around the world. UTIAS/SFL is a lab

consisting of students mentored and aided by experienced staff members in the the design, development and

building of small satellites. Over the years, the lab has developed the Canadian Advanced Nanospace eXper-

iment (CanX) projects as a platform to test research and development experiments in space. Collaborating

with other universities, as well as industry partners, the Space Flight Lab provides a quick and effective

way of preforming on-orbit experiments without long delays and high costs. The SFL satellites also have

the ability to quickly develop mission specific satellites outside the CanX platform for a variety of different

customers.

The UTIAS Space Flight Lab operates based on the mircospace philosophy, an approach to designing

nano- and microsatellites on quick schedules for relatively low cost. The Space Flight Lab can quickly

produce satellites due to the fact that commercial-off-the-shelf (COTS) products, not necessarily made for

1

Chapter 1. Introduction 2

use in space, are obtained at a low cost without long lead times. The applicable COTS solutions are then

tested to ensure functionality and mitigate any potential risk, making the need to only use radiation hardened

or spacegrade parts, unnecessary. Therefore, high cost and long lead times are avoided and allow for rapid

design and prototyping of new satellite technologies without delays. The idea of the mircospace philosophy

is to manage and overcome (calculated) risk as oppose to avoiding it entirely [1].

SFL Satellite Buses

There exist three different nanosatellite form factors that SFL is currently building. The satellite bus chosen

for a project depends on the specific mission and payload requirements. The first satellite is the triple

cubesat (3U) form factor, a 10× 10× 34 cm , 3.5 kg satellite bus that was successfully used for the CanX-2

satellite, seen in Figure 1.1.

Figure 1.1: CanX-2 Satellite (Courtesy SFL)

The 3U satellite bus is a variation of the 1U cubesat design, and provides more space and capability for

additional payloads. The second satellite bus is the Generic Nanosatellite Bus (GNB), a 20 cm cube, with

a 7 kg mass, seen in Figure 1.2a. The GNB was designed using heritage from the 3U form factor and is a

larger version of the 1U that still retains the cube shape. Lastly is the Nanosatellite for Earth Monitoring

and Observation (NEMO) bus, the largest SFL bus, which is capable of generating more power for larger

instrument payloads. The NEMO satellite measures 20× 20× 44 cm in dimension with a solar array area of

0.3m2 and weighs 15 kg, seen in Figure 1.2b.


(a) GNB (b) NEMO

Figure 1.2: GNB and NEMO Satellites (Courtesy SFL)

Operational Satellites

To date SFL is successfully operating four satellites in orbit. The Microvariability and Oscillations of STars

(MOST) astronomy mission was Canada’s first space telescope and the first space science microsatellite.

MOST uses a high-photometic-precision optical telescope to conduct long-duration stellar photometry in

space [2]. Following the 2003 launch of MOST, the CanX-2 satellite was launched in 2008. CanX-2 ues

a triple cubesat form factor and was host to several different experiments from contributing universities,

including a GPS radio occultation experiment, an atmospheric spectrometer, a network communications

experiment and a space materials experiment [3]. The CanX-2 satellite is highly successful and is still

operational and returning useful data. Also launched in 2008 on the same launch as CanX-2 was CanX-

6/NTS (Nanosatellite Tracking Ships). NTS was developed and launched in only 7 months, demonstrating

the rapid and responsive capabilities of small satellites [4]. NTS was the proven precursor to AISSat-1, a

nanosatellite dedicated to investigate the use of an Automatic Identification System (AIS) sensor to track

maritime ships for the Norwegian government [5].

SFL launches their satellites using the Nanosatellite Launch Service (NLS) on several different rockets as

a secondary payload, making it less expensive then a dedicated launch. The on-orbit satellites are operated

through the SFL ground station.

Future Missions

SFL is currently constructing several new projects in addition to the CanX-7 deorbiting demonstration

mission being presented in this thesis. The CanX-3 or BRIght Target Explorer (BRITE) will be launched in

the near future and will be used to make photometric observations of bright stars to examine for variability

amongst the stars [6]. The CanX-4/-5 mission will be two identical GNB satellites launched into the same


orbit that will demonstrate formation flying using a propulsion system tested on the CanX-2 satellite [7].

The NEMO-Aerosol Monitoring (NEMO-AM) will detect aerosol content in the atmosphere using the largest

SFL form factor to accommodate larger and more power hungry payloads [8]. As well, owing to the success

of the first AISSat, a second AIS satellite, AISSat-2, will be launched into orbit to meet the growing demand

for more AIS tracking data.

1.2 Space Debris Accumulation

Previously it was thought that due to the vastness of space, overcrowding would not become an issue in the

near future. There was no thought put into deorbiting or removing any spacecraft that were placed into

orbit. However, owing to required satellite operations in specific orbits, the Low Earth Orbit (LEO) and

Geostationary Earth Orbit (GEO), are quickly becoming very crowded. Not only are satellites adding to

the debris, but the upper stages of launch vehicles, any fragments that break off on-orbit spacecraft and the

human presence aboard the International Space Station, are all contributing to the debris accumulating in

space. If nothing is done about the accumulation, operational satellites in high density orbits will face an

increasingly hostile environment and the risk to future missions will be too great to freely launch satellites

into space without extensive justification of the mission.

As the Space Flight Lab has a number of different LEO missions on the horizon due to successful past

missions, deorbiting future spacecraft will be an issue for SFL. In order to not contribute to the space debris

accumulation, a deorbiting strategy is required for the nano- and microspacecraft designed and built at SFL.

1.3 Canadian Advanced Nanospace eXperiment 7

The CanX-7 satellite is being developed in direct response to the growing problem of space debris in the

highly dense LEO. The primary mission of the CanX-7 mission will be to demonstrate a deorbiting technology

suitable for a cubsat platform. The deorbiting technology will be a modular drag sail that can be integrated

to the outside of a satellite and therefore provide minimum intrusion to the internal electronics and payloads.

Figure 1.3 shows an image of what the deployed deorbit device will look like.

The CanX-7 satellite will utilize the highly capable and proven 3U form factor. It will use passive

thermal control and electronics currently used by the GNB satellite as several components used on the

CanX-2 satellite are no longer available. The satellite will be powered using a lithium ion battery with body

mounted triple-junction solar strings and will use one housekeeping on-board computer for command and

data handling. Communications prior to and post-sail deployment will be accomplished through the use of


an UHF receiver and S-band transmitter, using the existing SFL ground station. The attitude subsystem

will consist of a magnetometer and three magnetorquers. While attitude components are being included,

they are being used to monitor and perform experiments post-device deployment and for secondary payload

operations prior to device deployment, as the demonstration of the deorbiting device will be completely

passive.

Figure 1.3: CanX-7 Satellite with Deployed Drag Sails

The CanX-7 mission has three main mission goals to accomplish with on-orbit operations:

− Demonstrate a deorbit device for cubesats

− Validate post-deployment deorbit and attitude models

− Operate secondary payload provided by COM DEV Ltd

In addition to the primary deorbiting demonstration, the CanX-7 satellite will be monitored post-

deployment to observe if the predicted lifetimes are within approximately 5 to 10 % of the expected values.

In addition, the attitude of the CanX-7 satellite will be observed to confirm attitude models. Prior to the

deorbiting device deployment, the CanX-7 satellite will operate a secondary payload, demonstrating that

the deorbit technology can be an addition to a satellite without interrupting any payload operations. The

secondary payload being flown will make CanX-7 the first satellite to carry an aircraft Automatic Depen-


dent Surveillance - Broadcast (ADS-B) receiver, provided by COM DEV Ltd, and will detect ADS-B signals

from space with on-board signal processing. The technology will demonstrate whether space-based ADS-B

technology can be used to provide a source of continuous surveillance for commercial aviation. Providing con-

tinuous surveillance for aircraft will make the need for having to use predetermined paths and standardized

waypoints, for aircraft flying over large bodies of water and land areas without radar coverage, unnecessary.

This will allow for better flight planning and reduced emissions and fuel consumption of aircrafts [12].

1.4 Scope

This thesis was motivated by the growing space debris problem and describes the process of developing the

deorbiting device that will be used on the CanX-7 mission and the associated analysis. First, an overview of

the deorbiting technologies that exist will be described in Chapter 2. Second, the lifetime analysis performed

to size the chosen deorbiting technology is presented in Chapter 3 and the design evolution and prototyping

is presented in Chapter 4. Finally the intended test plan and various risk reduction deployment experiments

are documented in Chapter 5.

The deorbiting technology demonstrates a modular design intended to deorbit cubesat platforms. An

understanding of how long the deorbit period of a satellite utilizing the chosen drag device was needed to size

the technology. Various parameters affecting the on-orbit lifetime of a satellite were analyzed to determine

the performance of the drag device. Parameters such as the drag coefficient, launch year, solar activity,

atmospheric density, projected versus actual drag area, starting altitude and ballistic coefficient were looked

at to see how they affected lifetime. All parameters considered provided a complex picture of the length

a satellite will remain in orbit. In parallel to the deorbit analysis, the deorbiting technology was being

designed using heritage obtained from proven SFL technologies. With each design iteration, a prototype

was made to determine the feasibility of the concept and whether the design could be scaled to different

required sizes. In addition to designing and performing lifetime calculations, a test plan was formulated for

the drag device. The chosen technology needs to be tested according to SFL guidelines and the required

testing to be performed on the deorbiting units required forethought such that they can still be tested once

integrated onto a satellite, without disturbing other vital satellite components. Moreover, deployment tests

were preformed to mitigate any potential problems that may arise during launch, stowage and on-orbit

operations. The documented lifetime analyzes, design evolution and testing procedures of the device give a

sense of the difficult but very possible task of working within small satellite constraints.

Chapter 2

Background: Space Debris Cause and

Mitigation

The rapid accumulation of space debris from the beginning of human operations in space to the present

day will eventually change the future of spacecraft launches. Of particular concern for the UTIAS Space

Flight Lab is the effect space debris will have on small satellites. Launching a number of satellites to test

new technologies on a rapid schedule is what defines the small satellite world as unique over larger satellites

which cost orders of magnitude more to develop and may take upwards of a decade from the initial design

before being launched. If restrictions on launching satellites without well defined deorbiting strategies are

enforced, the number of small satellite missions will be directly effected and will hinder future research. The

following section will discuss why space debris is becoming a problem which cannot be resolved without

action and what steps can be taken to mitigate future debris creation. As well, current technologies that

can deorbit satellites will be discussed and whether or not they can be utilized on small satellites.

2.1 The Space Debris Problem

Space debris has become a growing threat for orbiting spacecraft. Space debris can be defined as all man-

made objects, including fragments, in Earth orbit or reentering the atmosphere that are nonfunctional [13].

Currently, more than 20 000 items are being tracked in space, which includes operational spacecraft in

addition to space debris above 10 cm [9]. Any object, whether it be a spacecraft fragment, is increasingly

difficult to monitor if it is smaller than 10 cm. However, a debris piece in the 1 cm−10 cm range is capable of

debilitating a satellite upon impact and therefore becomes very dangerous considering the inability of these

7

Chapter 2. Background: Space Debris Cause and Mitigation 8

objects to be properly tracked.

The debris problem has become a focus of attention in recent years due to the 2009 collision between the

defunct Cosmos satellite and an operational Iridium satellite [10], as well as the 2007 Chinese anti-satellite

demonstration [11]. Although satellites are being tracked, as well as their orbit positions propagated, it is

still very difficult to determine if a collision will occur due to the continually changing nature of the Earth

atmosphere. Both incidents left large clouds of debris that now pose a risk to current and future operational

spacecraft.

In order to try and prevent future collisions, the Inter-Agency Space Debris Coordination Committee

(IADC) is calling for the deorbit of nonoperational satellites within 25 years after the completion of their

mission, or within 30 years after launch if they cannot be reorbited to a graveyard orbit where the chance

of collision drops dramatically [13]. This time period was deemed sufficient for satellites to deorbit and as

a result not significantly contribute to the growing space debris problem. Left to deorbit naturally, small

satellites can take upwards of a 100 years to deorbit depending on initial altitude and ballistic coefficient,

which will only further aggravate the problem. A visualization of the current debris problem in LEO and

GEO can be seen in Figure 2.1.

Figure 2.1: Space Debris Visualization [14]

Allowing nonoperational satellites to remain in orbit increases the risk of collision with not only opera-

tional satellites, that are still providing useful data, but also other nonoperational satellites or debris. The

problem of space debris is therefore capable of growing exponentially and requires attention before drastic

measures need to be taken to make sure launching a spacecraft will not meet with a costly end.


The only method used to deorbit satellites in the past was to use extra fuel remaining from primary

mission operations. This was possible because the satellites were very large and adding an extra margin to

the fuel budget did not hinder the mission. Other than using a dedicated burn, a satellite would only reenter

within 25 years naturally if its starting altitude was low.

2.2 Significance of Deorbiting Satellites

The guidelines placed forth by the IADC are not yet law in Canada; however, not showing any deorbit

analyzes for a satellite will make obtaining certain licenses required for small satellites more difficult. This is

a concern for small satellites and SFL in particular due to the fact that the number of missions in the future

could be significantly reduced. Considering that the nano- and microsatellites that SFL produces are suited

for quick, responsive missions, it is in the best interest of the lab to ensure that future small satellites are

able to deorbit and not contribute to the space debris problem.

Moreover, sending satellites into space imparts a stewardship amongst the users to ensure future space

operations will continue. Leaving the space debris problem for future generations to solve is impractical and

irresponsible. Therefore, developing a deorbiting device that can be utilized by a small satellite to ensure it

will reenter within the given time frame will provide a responsible and convenient solution once the guidelines

provided by the IADC are enacted into Canadian law and must be followed.

2.3 Mitigation Methods

In an effort to reduce space debris several deorbiting strategies currently exist for satellites. The attached

deorbiting device would operate at the end of mission of the satellite and depending on the method and

initial orbit, would take weeks to years to deorbit a particular satellite. However, owing to the mass and

volume constraints of small satellites, not all methods and devices are necessarily applicable to nano- and

microsatellites. Discussed below are several methods that were considered when selecting a deorbit device

that would be suitable for SFL sized spacecraft.

2.3.1 Active Methods

A deorbiting device is considered an active method when the satellite is required to be operational and able

to send and receive data during the deorbit phase in order to function. For example, using a rocket to deorbit

a satellite at end of life would be considered an active method as the attitude of the satellite would need to

be controlled to ensure that the thrust was applied in the proper direction such that the satellite is deorbited


and not reorbited to a higher altitude. Other active methods include pressure maintained technologies or

solar sails. The former would require carrying pressurized gases on the satellite and may cause issues with

obtaining a launch as it poses an increased risk to other payloads. The latter also requires active attitude

control to properly point, as well as large areas to be used as a propulsion source. The solar sail, as seen

in Figure 2.2, is intended to be used for deep space missions as it uses the Sun as a constant never-ending

propulsion source, but can be utilized as drag sails in LEO due to the difficulty in using solar pressure to

overcome the dominant drag forces.

Owing to the problems associated with having an operational satellite in order to use an active method,

it was determined that the particular deorbit solution for SFL would be an nonactive method. This would

simplify the deorbit phase of the mission and would not require future missions flying the deorbit device to

extend their component operational lifetimes to be able to handle upwards of a 25 year deorbit. This would

not be possible as small satellites typically have a designed operational lifetime of 2 to 3 years.

Figure 2.2: Solar Sail [15]

2.3.2 Passive Methods

A deorbiting device is considered a passive method if at the end of life of the satellite, it need only activate the

device and from that point on, no longer be operational for the deorbit device to function. Several methods

exist to passively deorbit a satellite, and can be accommodated on a small satellite. These methods include

rigidizable inflatables that do not require a constant pressure source after deployment, electrodynamic tethers

and drag sails. Rigidizable inflatables, however, still require the use of a pressurized container to initially

deploy and again will cause problems with obtaining a launch for the satellite with this device onboard.

The devices utilize a large area to provide drag as can be demonstrated by the Gossamer Orbit Lowering

Device (GOLD), seen in Figure 2.3. Electrodynamic tethers use a current through a long tether to interact


with the Earth’s magnetic field to remove energy from the satellite’s orbit, causing it to deorbit. Tethers,

seen in Figure 2.4, are difficult to test in a 1g environment due to their length, which can be lengthened to

decrease deorbiting lifetime. However, increasing the length causes problems in itself, such as arcing, chance

of severing the tether in space increases and there also exist several deployment issues with current designs

[17]. Drag sails work on the principle of using a large area to interact with the Earth’s atmosphere in order

to provide drag that will decrease the ballistic coefficient of the satellite causing it to deorbit faster.

Figure 2.3: GOLD Rigidizable Device [16]

Figure 2.4: Electrodynamic Tether [17]


Currently, a number of small satellite deorbiting technologies are being developed in industry and research

settings. A number of tether projects are being implemented to show the tether concept can be used as a

viable deorbiting method. Tethers meant for different sizes of satellites are being developed, as well a

method for removing existing orbiting debris, by Tethers Unlimited, Inc. (TUI) [18]. In addition, drag sail

technologies are being tested at Clyde Space [19] as well as the Deorbitsail [20] in partnership with Surrey

Space and EADS Astrium among other partners. All current drag sail technologies being developed represent

a centralized device, where the entire sail is stowed in a single enclosure, attached to the outside of the host

satellite. Therefore, scaling the device to accommodate different sized satellites can run into problems if the

cross-sectional area must remain constant. In this situation the device must grow in depth and will end up

potentially interfering with other satellite components.

2.4 Deorbiting Device for CanX-7

Because of the need to test in a 1g environment, deorbit passively and the ability to adapt to different small

satellite sizes, the drag sail was chosen as the best device to demonstrate a deorbiting technology onboard

the CanX-7 satellite. The orbital lifetime of an LEO satellite is affected primarily by atmospheric drag and

to a lesser extent by solar radiation pressure and luni-solar perturbations [22]. If a satellite starts off in

a higher orbit than 800 or 900 km, the drag sail is limited in its deorbit capabilities as atmospheric drag

reduces with increasing altitude and there is a lack of aerostabilization for the device to properly function.

However, as SFL satellites operate in the LEO environment below 850 km, the drag sail was considered the

most suitable option after a trade study was performed [21] to compare the different options.

The drag sail device will be a modular design with the drag area spread over four identical units. One

drag sail device will be made up of four units, where more can be added for larger satellites if the units do

not interfere with one another during deployment and operation. Each device will be deployed individually

and only from a command sent from the SFL ground station. The drag sail device will be low in mass and

take up minimum volume when stowed so as not to interfere with the satellite.

The modular approach to the drag sail will allow for a more flexible drag device that can be better tailored

to the host satellite. Unlike the drag sail devices mentioned at the end of Section 2.3.2, a noncentralized

device can be used on a number of different satellite form factors without the need for extensive modifications.

Chapter 3

Mission Analysis

To properly design the drag sail device, an initial estimate of the required drag area was required. The

estimate was obtained using Satellite Tool Kit (STK) and running a number of simulations with different

orbits and satellite properties. This work was accomplished in parallel with choosing a drag sail design

so that it could be determined whether a particular design could provide the required area. This chapter

describes the model used to run the STK simulations, the satellite properties that were varied and analyzed

and the error approximations associated with the lifetime results. Also included is a brief look at the intended

on-orbit operations plan envisioned for the drag sail being demonstrated on the CanX-7 mission.

3.1 Lifetime Model

To obtain accurate results, an STK model was selected that represented the design space being considered for

the CanX-7 mission. The intent of the drag sail device is to deorbit a 3U satellite, as well as be upgraded for

use on future SFL satellites. As a result, the analysis presented focuses on the three classes of nanospacecraft

at SFL: 3U, GNB and NEMO. Performing the analysis for all three satellite buses will help understand how

sensitive the lifetime analysis is to different form factors and will be used as a starting off point for sizing the

drag area for future GNB and NEMO buses that use the drag sail. Doing the analysis for all the SFL buses

will also provide more data points for a more reliable analysis. The CanX-7 design space therefore consisted

of being able to size the sail drag area from a 3 to 15 kg bus, with orbit altitudes as high as 800 km, from

either a noon-midnight or a dawn-dusk sun synchronous orbit. While the sizing of the sail was done for the

larger buses, there still remains future work to be completed in order to integrate the device to the larger

satellite buses as well as more analysis when CanX-7 validates the models. The STK model used consisted

of the following parameters, with the High Precision Orbit Propagator (HPOP) propagator:

13

Chapter 3. Mission Analysis 14

− Area/mass ratio, with out an attached drag sail: 0.0057m2/kg (3U), 0.0057m2/kg (GNB), 0.0043m2/kg

(NEMO)

− Area/mass ratio, with a 2.0m2 drag sail: 0.5714m2/kg (3U), 0.2857m2/kg (GNB), 0.133m2/kg

(NEMO)

− Area/mass ratio, with a 4.0m2 drag sail: 1.14m2/kg (3U), 0.5714m2/kg (GNB), 0.267m2/kg (NEMO)

− Drag coefficient: 2.2

− Solar radiation pressure coefficient: 1.0

− Atmospheric density model: NRLMSISE 2000

− Flux/Ap file: SpaceWeather-v1.2

− Geomagnetic update rate: 3-hourly interpolation

− Geomagnetic flux source: read Ap from file

Where the area to mass ratios were calculated by assuming the attitude of the satellite is fixed. In the

case of the area to mass ratio without a sail, the average area of two sides of the satellite face (if different)

were taken, as the required attitude performance of the secondary payload of the CanX-7 satellite has not

been completely determined. The area to mass ratios for the 2.0m2 and 4.0m2 were determined by dividing

the drag area by the mass of each corresponding satellite mass. The choice of the drag coefficient will be

further elaborated on in Section 3.1.3 and the choice for the atmospheric density model will be explained

in Section 3.1.7. The other parameters listed were determined to give the most accurate results based on

deorbiting analyzes for the CanX-4 and CanX-5 satellites [24].

The STK Lifetime tool, which predicted the lifetime of the satellite being modeled at different orbits,

was used with the following parameters:

− Drag coefficient: 2.2

− Drag areas: 1.0m2, 2.0m2, 4.0m2

− Area exposed to the Sun: varied with drag area

− Mass: based on form factor

− Solar flux sigma level: 0

− Atmospheric density model: NRLMSISE 2000

− Orbits per Calculation: 10

− Gaussian Quadratures: 1

− Decay Altitude: 120 km [29]

− 2nd Order Oblateness Correction

− Rotating Atmosphere


− Solar File: SolFlx0909 Schatten.dat

The chosen drag areas modeled represent different effective drag areas which will be further explained

in Section 3.1.4. The chosen solar flux sigma level, orbits per calculation, Gaussian quadratures, 2nd Order

Oblateness Correction and the Rotating Atmosphere were chosen based on literature [28] and prior STK

work completed at SFL [24]. The decay altitude was set to 120 km as at this altitude an object will reenter

Earth’s atmosphere unless a significant amount of energy is added to reorbit [29]. The Solar file was chosen

to be the SolFlx0909 Schatten.dat and will be explained in Section 3.1.7.

The design space for the CanX-7 satellite consisted of the following orbits that were used in the STK

simulations:

− Orbit Altitude: varied from 500 km to 900 km, depending on analysis

− Sun Synchronous orbit with a varying inclination and Local Time of Ascending Node (LTAN) based

on altitude

This design space was chosen as the CanX-7 satellite will most likely be launched into an orbit described

by this parameter range. This is due to the fact that previous and upcoming SFL missions have been

or will be launched into these orbit parameter ranges and in order to start the analysis and get an idea

of the required drag areas, this design space was chosen. Once the CanX-7 launch orbit is known, the

analysis will be redone for the correct parameters and the analyzes presented herein will contribute to better

understanding the resulting lifetimes obtained.

3.1.1 Drag Sail Sizing

Based on initial calculations and STK simulations it was found that a drag area greater than 2.0m2 was

required to deorbit a 15 kg NEMO satellite from an 800 km altitude in 25 years. However, once attitude

calculations were done [26], which took into account that the drag area would not remain perpendicular

to the velocity vector for the entire deorbit period, there arose an issue of aerostability at high altitudes.

Therefore, an area greater than 4.0m2 with an effective area of 2.0m2 above altitudes of 600 km was required

to deorbit within the required 25 years, set by the IADC [13]. Section 3.1.4 will further elaborate on the

analysis of using 4.0m2 and different potential effective areas. To get a better understanding of how the

drag area affects lifetime, Figure 3.1 shows the lifetime required to deorbit a NEMO sized spacecraft from

several different altitudes, for several different drag areas.


0.1

1

10

100

1 1.5 2 2.5 3 3.5 4

Lifetim

e(Years)

Drag Area (m2)

800km, Cd=2.4 700km, Cd=2.4 600km, Cd=2.4 800km, Cd=2.2

700km, Cd=2.2 600km, Cd=2.2 25 Years

Figure 3.1: Required Drag Area at Different Altitudes

The graph shown in Figure 3.1 depicts a 4.0m2 drag area as being capable of deorbiting a NEMO bus as

the curves fall below the 25 year line. However, this is not the case, as will be evident through the analysis

presented in this chapter. There is a certain error associated with these calculations and this particular area

will not be sufficient to guarantee a deorbit within a 25 year lifetime. Two different drag coefficients were

used in order to determine if there was a significant difference, and this will be further discussed in Section

3.1.3. The curves in the graph representing 600 km and 700 km do not intersect the 25 year line over the

given drag areas as the drag areas are sufficient to deorbit within the required lifetime. The density of the

atmosphere increases as altitude decreases and therefore less drag area is required to deorbit within a 25 year

lifetime from lower altitudes.

3.1.2 Interpreting the Lifetime Results

Once an approximate drag area was calculated, further analyzes were completed in order to determine the

sensitivity of the different satellite and orbit parameters used. The analyzes were done for each of the satellite

buses, and by varying an individual parameter and observing the effect it had on lifetime, an approximate

error seen in these calculations and simulations was calculated [27]. The following analysis varying individual

parameters was done using STK: first, different drag coefficients were compared, second the effective drag

areas experienced during deorbit was varied, and finally the effects of the solar cycle were analyzed. The

analysis was done in order to understand the effects different factors have on lifetime and ensure that there

is sufficient margin such that the 25 year lifetime set by the IADC is met. Unless stated otherwise, all figures


use a drag area of 4.0m2 with an effective drag area of 2.0m2 above 600 km (explained further in Section

3.1.4) and use a drag coefficient of 2.2. The following sections below summarize the results.

3.1.3 Drag Coefficient

The drag coefficient varies depending on the shape of the object and the environment it is in and affects

the deorbit lifetime of a spacecraft. Once deployed the drag sail material will interact with the Earth’s

atmosphere at a particular altitude and deorbit the satellite it is attached to. The accepted value for the

drag coefficient stands at 2.2 [40] for most spacecraft; however, this value may not necessarily apply to small

satellites which don’t have the same volume and large solar array and antenna appendages. Several sources

mention utilizing a drag coefficient of 2.0 [23] as a value for smaller satellites and one calculation done at

SFL, based on collecting the on-orbit data from experiments done on the NTS satellite and comparing it to

the predictions from STK analysis, put the value as low as 1.28 [24].

0.1

1

10

100

900 800 700 600 500

Lifetim

e(Years)

Altitude (km)

Cd=2.2 Cd=2.0 Cd=1.28 25 Years

Figure 3.2: Comparison of Different Drag Coefficients

However, there are also papers that claim the drag coefficient increases with altitude [25], but again this

analysis was completed before small satellite development began. Therefore, several different drag coefficient

values were plotted in order to see the difference they have on lifetime. From this analysis a value of 2.2 was

utilized as the differences between the 2.0 and 2.2 coefficients does not vary significantly for the reference

spacecraft with a 15 kg mass at several different altitude values, seen in Figure 3.2. The graph also shows

that both the 2.2 and 2.0 drag coefficients both give lifetimes of less then 25 years for the 15 kg spacecraft

at 800 km. The 1.28 coefficient was found to be too unrealistic and the analysis showed a lifetime much


greater then 25 years. Selecting 1.28 without a review of the method used to calculate the coefficient and

without further references pertaining specifically to small satellites would make the analysis too conservative.

The kink seen in Figure 3.2 at 700 km is not easily explained and may just require more data points for a

smoother curve.

3.1.4 Effective Area

In altitudes higher than 600 km it has been found that in the current device configuration, aerostabilization

will not occur. Other disturbance forces dominate over the aerostabilizing forces [26]. Therefore, the drag

area provided by the drag sail device may not remain perpendicular to the velocity vector for the required

amount of time necessary to deorbit to an altitude where aerostable forces will dominate even if the centers

of mass and pressure of the satellite are placed appropriately for that configuration to occur. Below a certain

orbit altitude, approximately 600 km [26], the aerodynamic drag forces will dominate and the drag sail will

stay perpendicular to the velocity vector for the majority of one orbit. However, if the start altitude is large,

the drag sail will take a very long time to deorbit to 600 km and the lifetime will be greater than the allowed

25 years.

In an STK simulation, it was not possible to quickly determine what the attitude of the satellite would

be due to the dominate disturbance forces at a particular altitude. Therefore, it was difficult to know what

percentage of drag area would be available over one orbit. To account for this in a simulation, a percentage

of the available drag area was simulated from the start orbit to about 600 km and then it was assumed

all the available drag area would be utilized after 600 km. The effective area, as used in this thesis, can be

defined as the percentage of area that the drag sail is assumed to use to contribute to deorbiting the satellite.

In this way, the amount of time the drag sail does not spend perpendicular to the velocity vector due to

aerodynamic-instability, not helping the satellite to deorbit, is accounted for in the percentage of drag area

being simulated in the analysis. Therefore, from this analysis, the amount of effective area required for the

drag sail to deorbit a satellite within 25 years can be determined and from here attitude calculations can be

done to show the effective drag area will be greater than this.

Figures 3.3 - 3.5 demonstrate the simulated effective areas of 100%, 50% and 25% of a 4.0m2 drag area.

A drag coefficient of 2.2 was used for the three different satellite form factors at different altitudes. For

the 3U satellite bus analysis, shown in Figure 3.3, the lack of aerostabilization does not cause a significant

problem and the lifetimes for all effective area cases is below the required limit. This is most likely due to the

mass of the bus for the given drag area, more of which will be discussed further in Section 3.1.7, pertaining

to the ballistic coefficient.


0.01

0.1

1

10

100

800 700 600 500

Lifetim

e(Years)

Altitude (km)

100% Drag Area 50% Drag Area 25% Drag Area 25 Years

Figure 3.3: 3U Effective Area Lifetime

The effective area cases were also determined to not cause a problem on a 7 kg satellite bus, seen in

Figure 3.4. However, the 25% effective area case is quite close to the 25 year line at 800 km. Owing to the

potential error seen in the simulations, discussed further in Section 3.1.7, this might become an issue if a

GNB configuration satellite is launched into a high altitude with the drag sail device attached where around

25% effective area can be expected.

0.01

0.1

1

10

100

800 700 600 500

Lifetim

e(Years)

Altitude (km)


Figure 3.4: GNB Effective Area Lifetime

However, for the 15 kg satellite bus, having an effective area of less than 50% results in a lifetime greater


then 25 years at 800 km, seen in Figure 3.5. If the expected effective area is around 25% then drag sail will

not be sufficient to deorbit within the required time period. However, like the 25% case with the GNB bus,

the 50% effective area for the 15 kg bus with an added error margin might be just over 25 years. Therefore

the case of 50% effective area for the 15 kg bus is further analyzed in Section 3.1.7.

0.1

1

10

100

800 700 600 500

Lifetim

e(Years)

Altitude (km)


Figure 3.5: NEMO Effective Area Lifetime

A similar trend is seen across all three graphs in Figures 3.3 - 3.5. The data lines all converge to a single

line due to the fact that the calculated lifetime for the particular bus from 600 km to deorbit was added

to the lifetime from the start altitude to 600 km for all altitudes greater than 600 km. While this was the

approach taken, the method may result in a lifetime error. As will be discussed in Section 3.1.5, the lifetime

of a satellite also varies depending on the launch year. For the effective area simulations, if the lifetime for

a 15 kg bus at an initial altitude of 800 km was calculated, both the deorbit time from 800 to 600 km with

a percentage of the area used and the time from 600 km to 120 km with the full drag area being utilized,

were both calculated with the same start launch date. However, the second lifetime calculation of 600 km

to 120 km should really be done at the year when the satellite reaches 600 km. Take for example the 15 kg

satellite from a start orbit of 800 km, launched in 2013 with a 50% effective area, demonstated in Table

3.1. The lifetimes are different but do not vary by more than a couple of weeks and can be both over- or

underestimated. This potential error is accounted for in the overall error attributed to the STK calculations,

but as will be seen in Section 3.1.7, it is not the greatest source of error and will most likely not effect the

lifetime approximations significantly if at all.

Therefore, for both the 3U and GNB configurations, the drag sail will deorbit the satellite with 25 years


Table 3.1: Small Error in the Effective Area Lifetime Calculations Example

Year Lifetime (Years) Lifetime (Years) Total Lifetime

to 600 km from 600 km to 120 km (Years)

2013 21.7 0.64 22.34

2034 n/a 0.46 22.16

if less than 100% of the 4.0m2 drag area is seen. However, for the NEMO configuration, if less than a 50%

drag area, 2.0m2, is seen before 600 km, the satellite will take greater than 25 years to deorbit.

3.1.5 Solar Effects on Lifetime

The lifetime of a satellite can vary depending on when the satellite is launched with respect to the solar

cycle. The lifetime can vary by a couple of years depending if the satellite is launched in a solar maximum

or minimum and depends on how many solar cycles encompass the deorbit period. One solar cycle is

approximately 11 years and consists of a period of high and low solar activity. F10.7 values, explained

further in Section 3.1.7, are used as indicators to determine the solar activity by STK [29].A drag area of

4.0m2 with an effective area of 2.0m2 used above 600 km was used, with a drag coefficient of 2.2. Using

these drag area values and keeping all parameters constant while varying the launch date, several graphs

were obtained in Figures 3.6 - 3.8, that show the lifetimes when the launch date is varied.

0.01

0.10

1.00

10.00

100.00

Lifetim

e(Years)

Launch Year

800km 700km 600km 500km 25 Years

Figure 3.6: Lifetime for a 3U form factor versus Launch Year


0.01

0.10

1.00

10.00

100.00

Lifetim

e(Years)

Launch Year


Figure 3.7: Lifetime for a GNB form factor versus Launch Year

Both graphs in Figure 3.6 and 3.7 for the 3U and GNB satellites are under the required lifetime of

25 years. The graphs follow a pattern of repeating low and high lifetimes. The pattern roughly follows

the solar cycle, where periods of high solar activity give lower lifetimes and low solar activity gives longer

lifetimes. The predicted solar activity can be seen in Figure 3.10 and will be further discussed below.

0.01

0.10

1.00

10.00

100.00

Lifetim

e(Years)

Launch Year


Figure 3.8: Lifetime for a NEMO form factor versus Launch Year


The graphs seen in Figures 3.6 - 3.8 were calculated using an initial solar file. This older solar file was

first used as it was the default file in the STK Lifetime tool. However, an investigation into which solar file

should be used proved that the default solar file was too conservative, described further in Section 3.1.7.

Therefore, to make sure the lifetimes were below 25 years with the new solar file, the calculations were redone

for the 15 kg bus at 800 km. This is the worst case that will experience the highest lifetimes and was solely

looked at as opposed to updating all the data which would be laborious and unnecessary as the pattern for

the other cases would hold and only vary by a couple of years. The updated case can be seen in Figure 3.9

and shows that with the new solar file, the lifetime is below the 25 year mark.

1

10

100

Lifetim

e(Years)

Launch Year

NEMO, 800km 25 Years

Figure 3.9: Lifetime for a NEMO form factor versus Launch Year, Updated Solar File

The graphs show that the lifetime of a satellite with the attached drag sail device can vary by almost

4 years depending on the launch date. The graphs consist of launch dates from 2011 to 2030 to encompass

almost two solar cycles. However, in the graphs, the data collected for the GNB and NEMO satellites at an

altitude of 800 km does not follow the pattern of varying lifetime according to the solar cycle maxima and

minima. The 3U bus at 800 km also does not follow the pattern when compared with Figure 3.10 due to the

fact a solar minimum in the year 2016 should cause a higher lifetime, but this happens in the years 2013

and 2014. This seems to indicate a problem with the 800 km models for all three satellites. One possible

explanation for the discrepancy in the GNB and NEMO cases may be due to the fact that at 800 km the

lifetime for both satellites encompasses almost or more then two solar cycles. The variation seen at the lower

altitudes all fall within one solar cycle and the effects of the solar activity on the satellite may be felt more

acutely when a solar maximum occurs as oppose to a solar minimum, whereas this may not be so prominent

at higher altitudes where the Earth’s atmosphere is not as dense and there are fewer particles for the drag


sail to interact with. In the 3U case, the 800 km altitude lifetime may be slightly shifted due to the possibility

that the solar activity may have a slightly premature effect on small satellites. More analysis into why these

effects are seen at a 800 km altitude and why they affect different sizes of satellites buses is required to fully

comment on the results. The relevant information found from these graphs shows that even with the change

of launch year, 4.0m2 drag area with an effective area of 50% is able to deorbit all SFL satellites within the

required 25 years.

The graph shown in Figure 3.10 represents the variation of the solar flux over a number of years, which

represents the solar files used in STK with the lifetime tool. These models give a good representation of

overall solar activities during a given solar cycle. There are recorded historical data and predicted future

values seen in the graph. There is a certain error associated with obtaining the future values as they predict

values based on current information [28]. However, the length of a solar cycle can vary and is not necessarily

11 years and this graph may not be an accurate representation 25 years into the future. Though the potential

problems, it is a good approximation to use in predicting a lifetime based on a number of varying factors

given there is no guaranteed way to predict future solar activity levels accurately.

Figure 3.10: Solar Flux Predictions over Several Solar Cycles [28]

Figure 3.10 demonstrates that while the solar cycle is low the predicted lifetime is longer. This trend

is matched when the predicted lifetime graphs while varying launch year are compared to the predicted


solar cycle intensity. During the solar cycle, the density of Earth’s atmosphere changes with the increasing

or decreasing solar activity levels. When the solar cycle is high, the density increases and there is more

drag experienced on orbiting spacecraft. When the solar cycle is low, the density decreases and less drag

is experienced, increasing the lifetime. As the most significant amount of time is spent in high orbits, (see

Figure 3.11 for example of the varying altitude over the deorbit period for a NEMO sized spacecraft) if the

solar cycle is high during the launch period, the satellite will deorbit faster when compared to launching in

a period of low solar activity.

Figure 3.11: Orbit Altitude versus De-orbiting Time

As a result of the effect the solar cycle has on the lifetime of a satellite, it is important to take in account

the launch date for the CanX-7 satellite in order to get a more accurate result of the predicted time spent

deorbiting.

3.1.6 Time-Area Product

Further recommendations have been added to the IADC guidelines, from NASA, that the device chosen

to aid deorbiting reduces the time-area product or proves that the resulting on-orbit collision between the

deorbiting device and another object will not lead to the creation of new debris [30]. The time-area product of


a satellite is the time, in years, taken to deorbit to 120 km (the orbit altitude at which a satellite will deorbit

within a matter of days without the addition of significant amounts of energy to reorbit [29]) multiplied by

the drag area (m2).

The following graph, seen in Figure 3.12 was produced for a 15 kg satellite with a drag area of 4.0m2, an

effective area of 2.0m2 above 600 km, over altitudes of 900 km to 600 km, to compare the time-area product

for a satellite with and without a dedicated drag device.

1

10

100

1000

900 800 700 600

TimeArea

Prod

uct(m

2 years)

Orbit Altitude

NEMO with Drag Sail NEMO with out Drag Sail

Figure 3.12: Time Area Product for a NEMO Satellite with and without a Drag Device

However, as the graph in Figure 3.12 shows, the time-area product above 800 km is more or less the

same with or without an attached drag sail. It is only after 800 km that the drag sail starts to reduce the

time-area product. This difference may be due to the fact that atmospheric density starts to increase as

altitude decreases and the drag force becomes stronger and introduces a nonconservative force that takes

energy out of the satellite orbit. Therefore, the graph is saying that the drag sail is more effective in lower

altitudes. Since the time-area product is reduced, the increased drag sail area will not pose a collision risk

with other spacecraft or debris pieces because the time it takes to deorbit is reduced. In addition, the drag

sail is mostly made up of a thin plastic, discussed further in Chapter 4, and the resulting collision between

the primary area of the drag sail and a piece of debris will result in a low-energy collision when compared to

a collision between a debris piece and the satellite. A low-energy collision would result in damage to the sail

area, such as a hole or tear to the sail material, but will not result in the additional creation of debris that

would be harmful to other satellites. Therefore, although the drag sail does not significantly decrease the

time-area product, a resulting collision with the sail material will not result in a large debris cloud further

aggravating the growing debris problem.


The other satellite form factors had a similar trend as the NEMO bus for the time-area product with

and without the drag sail device. However, the 3U spacecraft shows a slightly larger difference in time-area

product with and without a drag sail, most likely due to the fact that it is lower in mass and the drag sail

reduces the ballistic coefficient, further discussed in Section 3.1.7, more for the 3U case than in the NEMO

case.

3.1.7 Error Approximations

Calculations of the lifetime for a deorbit device are approximations owing to the errors associated with

predicting the future of the solar cycle and the density of the atmosphere. Therefore, in order to ensure the

lifetime calculations are representative of the behavior seen on-orbit, an error analysis needs to be performed

to assess the impact each factor has on the overall lifetime. From here, the factors affecting the lifetime can

be ranked based on the impact it has on the lifetime and an error can be associated with it. The accepted

standard of 2.2 [40] was used for the drag coefficient in all the following analyzes. The factors considered to

have a large impact on the calculated lifetime of satellite include:

− Solar Variations

− Ballistic Coefficient of the satellite

− Density of Earth’s Atmosphere - Atmospheric Model selection

First, the variations of solar predictions can be demonstrated by comparing different predictions of solar

flux found in file options used with the STK lifetime tool, seen in Figure 3.13.

0

5

10

15

20

25

30

35

40

Lifetim

e(Years)

Solar File Name

Figure 3.13: Variations amongst STK Solar Flux Files


The analysis of solar variation effects was done on a 15 kg NEMO bus launched in 2013, with an 800 km

orbit altitude, a drag area of 2.0m2, a drag coefficient of 2.2, from an LTAN of 1100 and using the atmospheric

model NRLMSISE2000. All these parameters were held constant while varying the solar file in order to

determine how each solar file differed from another.

The average lifetime of the above analysis was found to be 22.8 years and the median was 24.2 years. The

lifetime used in sizing of the sail was found to be 22.8 years when using the SolFlx0909 Schatten.dat file.

Therefore, this new solar file was chosen over the previously used SolFlx Schatten.dat file. The change to a

new solar file was made after the analysis showed that using the SolFlx Schatten.dat file produced a lifetime

greater then that of the average of the other files. The Schatten solar flux files provided by STK contain

values of monthly mean 10.7 cm solar radiation flux (F10.7) and geomagnetic index (Ap), which predict the

solar activity for the analysis period desired. The SolFlx0909 Schatten.dat file was chosen due to its average

lifetime given when plotted and being the most suitable for the given analysis. The graph in Figure 3.13

depicts a wide variation of lifetime values when using different solar flux files. As it is very difficult to predict

the solar activity in the coming years it is similarly difficult to say which solar file should be used. Using

the most conservative value, predicting very low solar activity, would be impractical as it would oversize the

sail and lead to over constraining the design. However, using the high solar activity predictions would lead

to undersizing the sail and difficulty ensuring the satellite would deorbit within the given time frame.

Second, changing the ballistic coefficient of a satellite changes the way it interacts with the space environ-

ment and as a result will impact the lifetime spent in orbit. Investigating the different ballistic coefficients

for the different SFL satellite form factors can be seen in Figure 3.14. It shows that a ballistic coefficient of

less than 3.4 kg/m2is required to deorbit within the required 25 years.

The ballistic coefficient used here and in all the preceding lifetime calculations can be defined as follows

[29]:

Cb =M

CdA(3.1)

Where M is the mass (kg) of the system, Cd is the drag coefficient and A is the drag area (m2). The

ballistic coefficients plotted in Figure 3.14 come from the three existing SFL satellite form factors using a

2.0m2 drag area. The full drag area of 4.0m2 was not used in this case since it made the graph easier to

plot as the ballistic coefficient was held constant as opposed to decreasing after 600 km once the full drag

area can be assumed. The higher altitudes were more important in determining what ballistic coefficient

and the corresponding SFL satellites can be deorbited. Therefore, the analysis using the smaller drag area

holds and is slightly more conservative while still being realistic.


0.01

0.1

1

10

100

800 700 600 500

Lifetim

e(Years)

Altitude (km)

0.8kg/m^2 1.6kg/m^2 3.4kg/m^2 25 Years

Figure 3.14: Comparison of Different Ballistic Coefficients

Given that mass and drag area are two properties that can be controlled, the ballistic coefficient can be

changed in order to change the lifetime of the satellite. To decrease the lifetime, the mass of the satellite can

be decreased, or the drag area can be increased in an effort to reduce the ballistic coefficient. Lower ballistic

coefficients are required to deorbit quicker as the ballistic coefficient is a measure of how aerodynamic a body

is, with a higher ballistic coefficient meaning the body is more aerodynamic [29], encountering less drag and

therefore having a longer lifetime in orbit.

Finally, the density of the Earth’s atmosphere affects the lifetime of a satellite as the drag device interacts

directly with the changing atmosphere to deorbit. The change in density is partly due to the variation of

solar activity. However, trying to predict the changing density is difficult to do, like predicting future solar

activity. Trying to make a model of the Earth’s atmosphere also includes modeling the Earth’s magnetic

field and the gravity field among other properties. The difference in models provided in STK can be seen in

Figure 3.15. Several different models were developed in different years and used theoretical and/or on-orbit

recorded data to predict a varying amount of different properties are compared. The average lifetime of

the models is 22.51 years and the median is 22.6 years for a reference satellite at 800 km with a drag area

of 4.0m2, effective area of 2.0m2 and a drag coefficient of 2.2. The chosen atmospheric model was the

NRMLSISE2000 as it is the most up to date model and provides a lifetime of 22.5 years which is consistent

with the average and median of the others. As NRMLSISE2000 is a newer model, it utilizes data from

on-orbit satellites, whereas older models such as the Jacchia models utilize solely theoretical predictions to

calculate different parameters to predict on-orbit conditions.


As all the above analysis and error approximations show, it is very difficult to ensure a satellite will

deorbit within 25 years with a high level of confidence. Looking at the different factors affecting the lifetime,

it appears that the difficulty in predicting atmospheric and solar flux models into the future account for the

highest level of uncertainty in the lifetime calculations. While the other parameters also contribute, it is

these two models which provide the largest variation that accounts for the uncertainty in predicting satellite

lifetimes, with the solar activity predictions being more difficult to predict and affecting the density models.

0

5

10

15

20

25

30

Lifetim

e(Years)

Atmospheric Model

Figure 3.15: Comparison of Different Atmospheric Models

Owing to the higher ballistic coefficient and the uncertainty associated with the density and solar activity

predictions, a 10% error should be applied to the drag sail area to ensure a lifetime of 25 years is achieved.

With the reference spacecraft, used to compare different models, the effective drag area would then become

2.2m2. However, the NEMO satellite bus used as the reference spacecraft has a 0.3m2 solar panel. This

solar panel will be utilized as drag area and will result in an overall drag area of 2.15m2 for the NEMO

bus when accounting for the approximate 50% effective area seen above 600 km. Therefore, the drag area

required to deorbit within the required 25 years is slightly undersized by 0.05m2. This slight difference in

required drag area will not make a significant impact on lifetime considering using a 4.0m2 sail with an

effective area of 2.0m2 deorbits in just under 25 years. Therefore, there is no reason to oversize the sail

further as the NEMO reference spacecraft will be deorbited within 25 years and the other two form factors

will deorbit well within the 25 year lifetime. This being said, there is still further analysis needed to ensure

that the NEMO bus will actually deorbit within 25 years. The analysis done in this chapter makes sure that

the CanX-7 satellite, a 3U platform, is deorbited within time limits and uses the NEMO bus as an example.


However, the NEMO bus is deorbited very close to the 25 year mark and will need further investigation that

once actually used on future NEMO missions, the drag area will be sufficient.

3.1.8 Expected Lifetimes

After all the analysis was completed to predict different lifetimes over a number of varying parameters, the

lifetimes for each of the satellite buses was done with a 2013 launch date to see a numerical comparison of

the drag sail’s function, seen in Table 3.2. The values in the table are assuming a drag area of 4.0m2 with

an effective drag area of 2.0m2, with a 2.2 drag coefficient and the chosen solar file and atmospheric density

model.

Table 3.2: Expected Lifetimes for a 2013 Launch

Altitude (km) Lifetime (Years), Lifetime (Years), Lifetime (Years),

3U GNB NEMO

800 6.64 11.28 22.34

700 1.06 2.38 7.74

600 0.14 0.28 0.64

500 0.03 0.06 0.12

As expected, all the lifetimes are below the recommended 25 year lifetime set by the IADC. Depending

on the satellite and the starting altitude, the lifetime of the satellite varies from a couple of weeks to a couple

of decades. Owing to CanX-7 being a deorbiting demonstration mission, it is ideal to have a short lifetime

such that a change in altitude can be seen quickly, however, it is also necessary to start from a sufficiently

high altitude to prove that the drag device is working and compare the actual deorbit to what was predicted.

The orbit selection for CanX-7 will be further elaborated on in Section 3.2.1 The current drag sail size allows

for a fast deorbit at high altitudes and the functional redundancy built into the device, discussed in Chapter

4, allows the option to control the amount of drag area to allow for a slower deorbit.

The expected performance for the CanX-7 satellite can be seen in Figure 3.16. The graph shows the

expected lifetimes for drag coefficients of 2.0 and 2.2 for different drag areas. Figure 3.16 shows that even

for 25% effective area, the CanX-7 satellite will deorbit well within the 25 year limit.

The drag area provided by the drag sail is oversized for the CanX-7 mission. This is done because the

expected lifetime of the satellite may change due to unpredicted solar behaviour and the problem of not

being able to aerostabilize over 600 km, as the drag sail remaining perpendicular to the velocity vector will

not occur at higher altitudes.


0.1

1

10

100

1 1.5 2 2.5 3 3.5 4

Lifetim

e(Years)

Drag Area (m2)

800km, Cd=2.4 700km, Cd=2.4 600km, Cd=2.4800km, Cd=2.2 700km, Cd=2.2 600km, Cd=2.225 Years

Figure 3.16: Expected CanX-7 Lifetime for a 2013 Launch for Different Drag Areas

3.2 Drag Sail On-Orbit Operations Plan

On-orbit operation of the CanX-7 satellite will consist of commissioning the satellite and operating the

primary and secondary payloads. The commissioning phase for CanX-7 is similar to other SFL satellites

and will follow a similar operations plan as other GNB spacecraft which can be found in [31]. The drag

sail deployment phase however, is unique to the CanX-7 satellite and the intended procedure is documented

here.

3.2.1 Orbit Selection

As discussed above, a wide number of parameters affect the lifetime of a satellite. The altitude of the satellite

affects the lifetime as the atmospheric density varies depending on where the satellite is. As discussed in

Section 3.1.8, for this deorbiting demonstration, it is ideal to be able to deorbit for a long enough period to

ensure that the deorbit models are accurate within a percentage, but to also ensure that a noticeable change

in altitude can be seen in a practical amount of time to prove that the deorbit device is working.

The ideal orbit to demonstrate the deorbiting device would be an orbit altitude around 700 km. From

this altitude, the drag sail would deorbit the CanX-7 satellite within approximately a year. The predicted

deorbit models could be verified against the actual performance and the measured solar activity over the

deorbit period could be used to recalculate the predictions. Moreover, a one-year deorbit period would be


feasible in terms of an operations budget as the satellite can not be monitored over a much longer time frame

without significant costs adding up.

Not only is the altitude of the selected orbit important to demonstrate the deorbit, but so are parameters

like the type of orbit and inclination. For SFL satellites, a sun synchronous orbit will most likely be the

orbit launched into. These parameters are important for the deorbit demonstration as a drag sail will

behave differently depending on the orbit. However, while specifying the type of orbit CanX-7 would ideally

be launched into would be good for demonstration purposes, it is not necessarily possible to obtain a desired

launch. Therefore, it is necessary to calculate the predicted lifetimes at different altitudes as described in

Section 3.1.

3.2.2 Drag Sail Deployment

Following the secondary payload operation, the primary drag sail payload will be deployed. The drag sail

device consists of four individual units, which will be further described in Chapter 4. Each drag sail unit

will be deployed individually from a signal sent from the SFL ground station. The reason to deploy each sail

unit individually is to observe the dynamics and capture data from a sail unit before the entire drag sail is

deployed and there arises a potential problem with sending or receiving data to/from the satellite.

The expected plan for deploying the sails will be as follows:

1. Determine attitude of satellite

2. If rates are low enough, deploy one drag sail unit

3. Determine attitude of satellite

4. Take photo of deployed drag sail

5. Check any sensors monitoring sail deployment

6. Transmit data down to ground station

7. Repeat until entire device is deployed

This will be the operations plan to follow for the first three drag sail units. After each unit is deployed,

the attitude of the satellite will be determined, to gage the effects of sail deployment on the satellite. Once

the attitude rates are sufficiently low so that communication is capable with the satellite, another sail unit

will deployed. However, after the third unit is deployed, a longer period will follow before the final unit

deploys to determine the attitude of the sail in a three-sail unit deployed configuration. There is a chance

that not all sail units will deploy, or it may be possible that while on-orbit, one sail unit may be damaged

and the drag area no longer usable. Therefore, it is worthwhile testing a three-sail configuration to determine

what attitude the satellite would settle into with only three drag sails deployed. The time taken between


each step in the drag sail deployment sequence will vary depending on how many passes it takes to transmit

and receive the required data and to determine the attitude of the satellite where necessary.

Step four of the deployment plan will have the satellite take a picture of the drag sail. Taking a photo

will quickly determine if the drag sail unit has deployed, as opposed to having to wait for a noticeable

change in the altitude to prove that deorbiting is taking place. The camera also provides the possibility of

photographing the drag sail well into the deorbit to determine if the sail material is damaged. Looking at

Two Line Elements (TLEs) provided by NORAD to track the altitude change will not determine if all four

of the drag sails have deployed, or the state of the sail post deployment. An alternative to using TLEs or

a camera to determine the drag sail has deployed, is to use an Earth based telescope to view the drag sail

and determine if all the sails have deployed. The sail material will be aluminized, as discussed in Section

4.3.3, and will therefore be reflective while in the sunlight. Therefore, it is possible, if the drag sail were

at a low enough altitude to be able to photograph the deployed device from Earth. Further calculations

and investigations into obtaining time and the cost associated with using a telescope remains to be done,

however, it does remain a possibility.

3.2.3 On-Orbit Experiments

On subsequent missions flying the drag sail payloads, the drag sails will be deployed at end of life of the

satellite, and no further satellite function is necessary. However, due to CanX-7 being a demonstration

mission, some experiments will be done prior, during and post sail deployment. Several experiments that

will be performed on the CanX-7 satellite, pending feasibility include:

− Evaluate the deorbit predictions

− Attitude Experiments

− Extent of communication with satellite while the drag sail is deployed

In order to evaluate the deorbit predictions, TLEs for the CanX-7 satellite need to be obtained at regular

intervals and compared to the expected performance for the specific launch date and properties of the satellite

and the orbit it is inserted into. Depending on the initial orbit altitude, it may not be practical to look for

a change in altitude until a couple of months into the deorbit. This will be determined once the launch for

CanX-7 has been arranged.

The attitude experiments will consist of determining whether the drag sail remains perpendicular to the

velocity vector to see how much drag area is actually being utilized. This will also factor into evaluating the

deorbit predictions since an effective drag area is only being estimated at this point.


The drag sail will cause problems with communications as it is highly reflective and quite large compared

to the satellite. The extent of communication can be estimated with an RF analysis, however, this can only

be confirmed on-orbit. There is also a potential concern of the reflective sail material interfering with the

individual drag sail deployment as each deployment would decrease the ability to communicate. However,

while this is a valid concern, it is mitigated by on-orbit perturbations that will change the attitude of

the satellite and the placement of the drag sails on the satellite to interfere as little as possible with the

communications antennas, shown in Section 4.7. While it may take more than one pass to deploy the final

sail unit, it is highly likely that a signal can be received.

Chapter 4

Drag Sail Device

The drag sail device will be the primary payload being flown on the CanX-7 mission. The CanX-7 satellite

will be a triple cubesat (3U) form factor. The 3U form factor was therefore the starting point for determining

the maximum cross-sectional area that the drag sail could be designed to, in order to fit on a cubesat face.

The drag device will consist of a sail material that will interact with the Earth’s atmosphere, in order to

deorbit, and will be held in place by a frame. The drag sail device is modular and will consist of four units.

As a drag sail is a new technology for SFL, and for cubesats around the world, ideas for the design of the

device were taken from several SFL deployables, from the CanX-2 satellite to the eXoadaptable PyrOless

Deployer (XPOD), a proven satellite separation system. Implementations of solar sails were also looked

into for inspiration for the drag sail design, as solar sails require deploying large areas with low mass that

must survive for long periods of time in outer space. This chapter describes the initial thought process

and development that went into designing the drag sail. It describes the prototyping, design choices made

and testing that went into choosing a design that would work with the CanX-7 satellite and that could be

upgraded to work with all SFL satellites in the future.

4.1 Initial Drag Sail Concept

Owing to the varying satellite structures at SFL, determining the design for a drag sail device was not a

trivial task. Initial drag sail concepts focused on devices that could be easily adaptable to all satellites as

well as scale with different sizes as was necessary. After several trade studies and discussions, it was decided

to make the drag sail device modular over a centralized design concept. While a requirement to be modular

was not necessary, there were advantages to choosing a modular design. With a modular approach, different

mounting placements could be accommodated on different buses, as well as slight adjustments to the size

36

Chapter 4. Drag Sail Device 37

for larger buses. With a centralized device, scaling the device would be difficult as would the placement on

all of the buses. The initial concept for the drag sail device can be seen in Figure 4.1, where the entire drag

area is separated into 4 equal circular areas.

Figure 4.1: Initial Drag Sail Concept (Courtesy of SFL)

The drag device was envisioned to be as simple as possible. The drag sail will consist of a sail material

held in place by a frame. The frame of the device will be stowed in tension and will utilize that energy in

deployment to unfold the sail material. The mechanism holding the device from deploying will be released

by a signal sent from the SFL ground station at the end of the satellite mission and will not require any

additional satellite power to function. The sections below will describe the constraints and requirements

placed on the drag sail device and why a modular approach works well for the CanX-7 mission and future

evolutions of the drag sail device to deorbit larger buses.

4.1.1 Deorbit Capabilities

The drag sail device needs to accomplish deorbiting within a period of 25 years after the end of life or 30 years

after launch. The choice of 25 years to deorbit was set by the IADC, whose guidelines are being implemented

by the Canadian government making it difficult to obtain certain licenses without showing that the satellite

in question will deorbit within the specified time frame in an effort to reduce debris. This could quickly

become a problem to the number of future small satellite missions which typically are built and flown using

both a relatively fast and low-cost approach.


In addition to deorbiting within a specific time frame, the drag sail is constrained by a number of different

parameters given that it is being designed to integrate with a small satellite. The CanX-7 satellite will be a

3U form factor and the drag sail will be mounted to one of the small faces of the satellite. This constrains

the maximum cross-sectional area of the device to be within 10×10 cm as will be discussed further in Section

4.1.2. Additional volume is taken up by the sensors that will gather telemetry about the deployment and

safe stowage of the device. Moreover, as the CanX-7 satellite will be around 3.5 kg in mass, the drag sail

must be less than 0.5 kg such that the 3U XPOD mass limits are not exceeded.

Particular attention to the materials chosen for drag sail components must be ensured due to several

requirements imposed on the device. The drag sail maximum lifetime is 25 years, therefore, the deployed

drag sail must survive up to 25 years or up to an altitude of 300 km (TBC). The drag sail will experience

high loads at this altitude that the sail is not required to survive as the satellite will come down within

days from this altitude without the need of a drag device. In addition to surviving the space environment

when deployed, the sail must survive stowage for a period of up to 3 years so that deployment is able to

occur. As well, to ensure the above is met, the drag sail will undergo extensive qualification and subsequent

acceptance testing, described in Section 5. These tests will occur in a 1g environment and therefore, the

drag sail materials need to be chosen such that they can survive repeated deployments in a 1g environment,

preferably without the use of extensive GSE.

In spite of all the requirements and constraints specified for a drag sail, it was chosen because it was the

best drag device to deorbit small satellites. It not only is a passive method that can be used at the end of

life of a satellite, but it can be made to fit on a variety of satellite form factors and is the right fit for SFL

satellites.

4.1.2 Dimensions

The initial dimensions of the deorbit device were constrained by the smallest face of a 3U bus. The dimensions

of the drag sail measured at 10×10×1 cm (width × length × height). The height of the device was constrained

at 1 cm due to the 3U XPOD height and the space restrictions on the bus itself, based on the CanX-2 satellite

which was a 3U bus. Initial lifetime analysis was performed for a reference NEMO satellite and a drag area

of 2.0m2 was found to be required. The drag area was divided into four separate units with a corresponding

drag area of 0.5m2. Each drag sail unit consists of a frame supporting the sail material that would interact

with the Earth’s atmosphere and cause the satellite to deorbit. The shape of each unit was initially chosen

to be circular to provide structural integrity to the design, shown in Figure 4.1.

As design of the drag sail progressed and consideration was given to the integration of the device to the


satellite, further space restrictions arose. The initial area of the drag sail was reduced due to the cutouts

required for the satellite legs, that interact with the XPOD. The satellite legs could not be removed for

the drag sails as they take launch loads and the drag sails are not required to take such large loads and

would over constrain the design. Therefore, cutouts were made in the drag sail housing to accommodate the

satellite legs, seen in Figure 4.2.

Figure 4.2: Maximum Drag Sail Cross Sectional Area

The required cutouts removed an area of 1 × 1 cm from the four corners of the drag sail housing. This

was a challenge to overcome as the cutouts made it more difficult to store a 0.5m2 drag area in an oddly

shaped volume. The problem was overcome with the final chosen method of stowing the frames used in the

device, the evolution of which will be described in Section 4.2.

4.2 Design Evolution

Several prototypes were made to test different concepts for the drag sail device based on the initial problem

of stowing the required drag area into a volume that would fit on a small satellite. The prototypes were made

using commercial off the shelf (COTS) materials to be able to quickly build and determine if the concept

was suitable. Each folding method for the frame of the device is described along with the initial method

used for folding of the sail material. At this point in time, each drag sail unit was circular. Once the best

method was found from analysis and prototyping, the circular method was changed and will be described in


Section 4.3.

4.2.1 Folded Frame

The first concept for the drag sail frame involved bending a carbon steel spring tape to fit within the required

area. The folded frame evolved as a concept in an effort to obtain the necessary stored tension to be able

to deploy the frames and attached sail material. The carbon steel spring tape was purchased as a COTS

material that is the treated raw material used for carpenter measuring tapes. The tape can store tension

energy when folded and was obtained with ease and at a low cost and proved to be quite useful in prototypes.

Below in Figure 4.3 is an image depicting the storage concept and its intended method of deployment.

Figure 4.3: Folded Frame Concept

The sail material would be stored on top of the folded frame, most likely with a thin divider so it would

not interfere with the frame deployment. This concept provided sufficient energy for deployment, however,

the frame fold was dependent on the bend radius of the frame material. If the frame had a large bend radius,

it was difficult to fit the frame within the required volume. This also made it difficult to scale the drag area

if the initial drag area of 2.0m2 was not enough, a problem which did arise after attitude calculations were

completed. A segmented frame was proposed in order to alleviate the dependence on the bend radius of

the material but did not have sufficient stored tension or structural stiffness when deployed. The segmented

frame also posed a problem as to how to connect the individual segments that could be stowed and deployed

and survive in space for upwards of 25 years.

4.2.2 Coiled Frame

The next idea, the coiled frame concept, tried to take advantage of the inherent properties of the tape

measure blade being used as the frame material. Instead of folding the material it was coiled around a spool

to store tension energy, as it is in a carpenters measuring tape. Except instead of adding another spring to


make the tape want to recoil, no spring was added such that the wound tape would want to unspool. Below

in Figure 4.4 is an image depicting the coiled frame concept and its intended deployment.

Figure 4.4: Coiled Frame Concept

The frame was coiled around a spool and took up as much of the area as it could, while the sail material

sat on top of the coil. Owing to the height constraint on the drag device, of 1 cm, the availability of COTS

solutions for the frame material was limited. Therefore, the coiled frame method of stowing and deploying

the drag sail appeared to be lacking the necessary stored tension. There also existed problems with choosing

a fold for the sail material that would be able to unfold while the device was deploying. Modifications to

this concept to suit the problem at hand would soon solve the problems seen with prototyping this concept

and will be discussed in Section 4.3.

4.2.3 Double Coiled Frame

To utilize the properties of the tape measure blade while still obtaining the necessary tension required, a

double coiled frame approach was considered. The double coiled frame involved having two frames coiled

about individual spools, where smaller tighter coils would allow for more stored tension energy. The method

also allowed for a linear deployment which made the selection of the material fold more straightforward.

Below in Figure 4.5 is an image depicting the storage concept.

Figure 4.5: Double Coiled Frame Concept

Several problems arose from the concept. Owing to the second spool required for the concept, the volume


was quickly used up leaving little room for the other components of the device. There also arose the issue

of how the two frames would be coiled individually and connect once deployed to close the circle and still

be able to draw the sail out from this attachment point, depicted in Figure 4.5 by the vertical arrow, so the

deployment remained linear and simple.

After experimenting with the several different stowage and deployment concepts, the shape of the indi-

vidual unit for the drag device changed allowing for the single coiled frame concept to be looked at in more

depth and is discussed in Section 4.3.

4.2.4 Initial Drag Sail Folding

Several different sail material folds were looked at in order to choose the best one that would be used for the

drag sail device. The fold of the drag sail material was important as it determined how the drag sail would

deploy and be stowed with the frame. It was important to ensure the material did not prevent the frame

from unfolding which may result in the material tearing or the device jamming. To determine the type of

fold to be used with the frame, several folds were investigated. Some criteria used to compare the folds can

be summarized as follows:

− Fold lines are perpendicular to tension lines during deployment (to aid deployment)

− Linear/non-centralized deployment (does not require more than two points to pull on to deploy the

stowed material)

− Ease of deployment with frame

− Fold area, dimension, thickness, pattern ratio (unfolded/folded area ratio)

− Strength of membrane after folding

− Number of creases formed during folding, storage and deployment (a lot of folds can damage the sail

material and affect its performance)

− Ability to scale to different sizes

− Ease/method of fold

Folds used in solar sails and folds that arose from nature, flower and leaf patterns, were used as starting

points to determine what fold would be best for the drag sail. Several of the criteria mentioned could not be

determined based on a quantitative analysis but were still important for determining how the material would

fold, unfold and hold up over the deorbiting timeframe. These properties were tested through prototyping

with the frame concept. Images of several of the folds are presented in Figures 4.6 - 4.9. The summary of

measurable dimensions and properties of each tested fold are listed in Table 4.1.


Figure 4.6: Butterfly Fold [32]

Figure 4.7: Leaf Fold [32]

Figure 4.8: Circular Fold [33]


Figure 4.9: Miura-Ori [34], [35]

Table 4.1: Summary of Calculated Properties of Different Folds

Fold Type Folded Folded Area Number Overall Number of Pattern

Dimension (m2) of Folds Fold Fold Cross Ratio

(cm × cm) Thickness Overs (Circle)

(mm)

Miura-Ori, 10 6× 6 36 10 2 16 8.2:1

Miura-Ori, 22.5 4 × 8.7 34.8 9 1.5 17 8.5:1

Miura-Ori, 45 2.5 × 17.1 42.75 11 1 24 8:1

Miura-Ori, 0 4 × 4.1 16.4 8 2 16 18:1

Circular n/a n/a 15 n/a n/a n/a

Tree Leaf (sq) 10.6 × 3.1 32.86 28 3 25 13.6:1

Tree Leaf 10.5 × 3.1 32.55 20 3 17 10.2:1

Butterfly (sq) 2 × 21.3 42.6 10 4 49 10.5:1

To determine the proper fold, the table was analyzed in detail. From the table, the Miura-Ori was

determined to be the best fold. Several angles, from 0◦ to 45◦ of the fold were tested to see if they differed

significantly. The angle refers to the vertical fold line angles. If the angle is 0◦, then the fold lines are

perpendicular to the horizontal fold lines. If the angle is greater, they are slightly offset, as seen in the left

hand image of Figure 4.9. A greater angle allows for an easier deployment, however, takes up more space

when folded. The largest angle that was tested was a 45◦ as folding any larger angle becomes increasingly

difficult and impractical when time is considered.

The several parameters that were considered important include the folded dimension, area, the number

of folds and the fold thickness and cross-overs. Also significant to choosing the best fold, nonquantitative

criteria needed to be tested to determine if the fold could deploy easily with the frame and if it could scale

to different spacecraft without changing the fold too drastically. Several fold prototypes were folded on


paper and Mylar, a representative sail material, to determine the fold handling criteria. Paper was used to

determine what the easiest way to fold the pattern would be before moving on to the Mylar material, as

it is more tricky to fold with. The Miura-Ori fold was the most versatile fold that could adapt to different

shapes and sizes. It was also the most reliable fold to both unfold and refold. The Miura-Ori was utilized

during prototyping of the drag sail to determine if a concept was feasible.

4.3 Updated Drag Sail Design

Throughout the design and prototyping of the original circular concept for the drag sail, it became apparent

that not enough structural stiffness would be obtained from the circular design as it was necessary to increase

the drag area due to updated attitude predictions. As well, stowing and deploying a circular device proved

to be more complicated. The initial circular design idea can be seen in a SolidEdge model below in Figure

4.10. The Figure shows the large overlap in drag area. Due to this overlap and other problems, the design

was updated and takes advantage of the coiled frame concept.

Figure 4.10: Original Drag Sail Concept

The new design concept has two frames, that will from here be referred to as booms, that both coil

around one axis on a spool with the drag sail material folded and stowed separately in front of the spool.

The drag sail device is no longer circular in shape, but results in a square with all the individual units being

trapezoidal in shape, essentially triangular in appearance. The units of the drag sail can be stacked on the


3U bus for CanX-7 or tiled for future use on larger buses. The specific details describing the changes of the

drag sail design, including the sail material, will be outlined in the next sections.

4.3.1 Design Overview

The drag sail dimensions were based off the required drag area to deorbit within the required 25 years

divided amongst 4 units. The new calculated drag area, that accounts for the sail not aerostabilizing at

high altitudes, is 4.0m2 which gives each unit a 1.0m2 area. To determine the dimensions of the drag area

and the angle at which the booms would deploy, a simple calculation was done. The drag area would be

trapezoidal, as the base of the drag sail, closest to the module, is not a sharp point. The area of a trapezoid

is defined as:

A =a + b

2h (4.1)

Where a is the length (m) of the top side, b is the length (m) of the bottom side and h is the height (m)

of the trapezoid. The length of b is known to be 0.07m, the drag area is 1.0m2, and h can be rewritten in

terms of b and the angle from which the booms deploy from the drag sail. Inputting the variables into the

equation, and varying the angle of boom deployment, the dimensions of the drag sail can be determined. It

was found that an angle of 45◦ for boom deployment was ideal and the dimensions seen in Figure 4.11 were

obtained, where the 1.4m dimension is the length of the deployed booms.

Figure 4.11: Drag Sail Dimensions


All four units will work together to provide a drag area in a square configuration when placed on the

satellite. Figures 4.12 and 4.13 show the drag sails stacked and tiled.

Figure 4.12: Drag Sail Device Placement on a 3U Satellite

Figure 4.13: Drag Sail Device Placement on a GNB Satellite

The modularity of the device allows for different placements on different bus configurations, where the


units do not necessarily have to be side by side to function. The drag sail material will be attached at three

points. Two attachment points will be at the tips of each boom, and the last attachment point will be at

the base of the module. In this way when the booms deploy they will also pull out the sail material that

will remain permanently fixed at the base.

4.3.2 Housing

The housing of the drag sail is the structure that stows the booms and sail material, as well as necessary

electronics and sensors, and protects those components from the harsh space environment. The drag sail

housing was originally planned to be made out of aluminum, but because it is one of the heaviest components

of the drag sail and quickly grew very complex to machine, the material was changed to a Windform material

[36]. The initial SolidEdge model for the updated drag sail concept is shown in Figure 4.14.

Figure 4.14: Initial Drag Sail Model

The spool in the model is oversized in order to determine if the sail material can be stored on top of

the spool. From testing it was shown that the spool should be smaller to coil the booms tighter and obtain

more stored energy and to separate the sail material completely from the booms. Also seen in Figure 4.14,

the guides at the corners of the opening of the housing to guide the booms out are not at the correct angles.

The angle was chosen such that the boom on the outside of the coil would have a straight deployment path


and the inside boom would have as close to a straight deployment path. However, it was determined that

in this configuration, the boom lengths would not be the same and woould end up causing problems and

both angles were therefore set at 45◦. The deployment path of the booms continued to be a problem during

prototyping as will be mentioned in Sections 4.4 and 4.6.

Separately stowing the material meant that it would be located at the opening of the housing such that

it was ready to be pulled out as the booms were deploying. The tips of the booms would be right beside

the opening and the rest of the booms would be coiled around the spool. However, this brought up the

problem that the sail material would be exposed to the space environment and could potentially fall out of

the housing if it were not properly restrained. Therefore, a door was added in a subsequent iteration of the

drag sail design, along with the change of the spool diameter and location, seen in Figure 4.15.

Figure 4.15: Old Drag Sail Concept with Door

The door was envisioned to aid with the sail material deployment. The sail material would sit on top of

the door, seen in Figure 4.16, which could have been designed with a raised edge portion so that the material

would sit inside. Once the drag sail was told to deploy the door would open and slide the sail material out

and the booms would release and unfold the sail. The door was designed after the deployable antennas on

the CanX-2 satellite [37]. The hinge was spring loaded, a closeup of the door can be seen in Figure 4.17,

and after a mechanism holding it in place was cut, the spring would open to its relaxed position and hold

the door away from the deploying booms and sail material.


There arose several issues with having a door mechanism in the drag sail design. First, the door added

significant mass to the overall device. Although the mass could be reduced through making cutouts in the

door, there was still additional mass in the system. Second, the addition of a door added another deployable

to the device that could cause problems. If the door of the device did not open, the booms could not pull

the sail material out and one unit or more would end up not deploying properly.

Figure 4.16: Door Tray

This added risk to the overall drag sail. Lastly, the presence of a door in the design as depicted in Figure

4.15 opening sideways, would cause timing issues during deployment. If the booms were to deploy before the

door was fully opened, they may get jammed in the device or may prematurely pull the sail out and cause

potential tearing of the material.

In the end the door, as designed, proved too risky to be included. However, the drag sail still needed

something to hold it in place during launch and stowage prior to deployment so that it would not fall out

of the housing. In order to constrain the sail material, it was proposed to have the boom material slightly

oversized such that the extra length could be bent in front of the housing opening and prevent the sail

from escaping. Once deployed, the extra length would not cause a problem to the operational drag sail and

would also not pose a problem during deployment. However, after the device was subsequently updated, as

described in Section 4.6, the door was added back into the design as it was determined to be necessary and

outweighed the risk.


The height of the module also posed a problem with finding available COTS products to use in the design.

The original height of 1 cm was still being used in the design and was proving to be a challenge. The only

available COTS tape springs were low in stiffness and could not store enough tension to deploy, let along pull

the sail material out. More on the boom material selection will be discussed in Section 4.3.4. It was at this

point that it was determined that the drag sail needed extra height as it was quite difficult to design within

the given constraints without using custom designed booms which would drive up the cost of the prototype

significantly.

Another observed issue with this prototype was the path the booms followed when deploying. Too much

volume was taken away by routing the booms directly to the device corners. Moreover, the little pegs at

the end were not strong enough to guide the booms to deploy at the required 45◦ angle. This caused some

concern as not having a straight boom path from the reel meant that the booms would deploy with additional

friction. However, too much room was taken away from the sail material stowage area.

Figure 4.17: Close-up of Door Hinge

Taking into account the all the above mentioned problems of the boom deployment path, the reel size,

the door issues and housing height, another design iteration was completed, seen in Figures 4.18- 4.20. This

design consisted of an increased module height from 1 cm to 1.6 cm that allowed the use of a COTS tape

measure blade that was capable of storing enough tension. The increase in height came from shortening the

satellite launch legs on the opposite side of the drag sails allowing for a total height of 6 cm for all 4 drag


sail modules stacked.

In each design iteration, the top covering of the module is not shown as it is not required. The intent is

to have the top plate of the satellite structure be the drag sail top mounting, so that additional mass is not

added to the system.

Figure 4.18: Drag Sail Housing

Figure 4.19: Drag Device Reel


The spool of the updated design was also modified to account for the larger tape spring height and the

additional length of the booms from the increase in drag area from 2.0m2 to 4.0m2. The groove on inside

of the spool, shown in Figure 4.19, is where the tape springs will be permanently attached, so that during

deployment they stop when the sail material is fully unfolded.

Giving the booms a path to deploy allows more stowage area for the sail material. However, routing

the booms through a particular path causes problems during deployment as there is a lot more friction

introduced. The problem of friction will be addressed in Section 4.4 discussing the prototyping of the drag

sail.

Figure 4.20: Drag Sail Device Housing

4.3.3 Sail Material

The drag sail material was chosen to be 12.7 micron double sided aluminized Kapton. A wide variety of

materials were considered as candidate choices for the sail material. During on-orbit operations, the sail

material faces many environmental threats, some of which include:

− Photon radiation (UV exposure)

− Charged particle radiation (ions, plasma)

− Extreme temperature cycles

− Micrometeoroid and debris impacts

− Photochemical deposition of contaminants


− Low Earth orbit atomic oxygen (AO)

− Vacuum

− Electrostatic charging

The extent to which each environmental threat affects the sail material is not only dependent on the

particular material, but also on the satellite and orbit parameters. The mission duration, the launch time

determining what solar activity the satellite is launched into, random solar events, and orientation of the

material with respect to the velocity vector among other variables will play a role in how the material will

fare in the outer space environment. A trade study of several potential materials, shown in Table 4.2 was

completed to determine the best material.

Table 4.2: Sail Material Trade Study

Trade/Option Aluminized Aluminized Aluminized CP1 [39] Gold Teflon

Mylar Kapton Kapton with Covered FEP

an Acrylic Kapton [38]

Coating [38]

Total Cost/Sail 57.70 USD 363.64 USD 180 USD 250 USD n/a 5

($/Sail)

Per-Roll Cost 750 USD 4000 USD 7495 USD 44586 USD 15000 USD 1500

($)

Lead Time 2 4 Based on n/a Based on 4

(Weeks) Availability Availability

UV Susceptibility No No No No No No

AO Susceptibility No No No No No n/a

RF Transparency No No No Yes No Yes

Space Heritage Yes Yes Yes Yes Yes Yes

Humidity May cause May cause No No No No

Susceptibility aluminum aluminum

to flake, to flake,

exposing exposing

Mylar Kapton

Due to the high cost and potential ITAR restrictions, aluminized Kapton with acrylic coating, CP1, and

gold covered Kapton have not been chosen as the sail material. Teflon FEP was also considered an attractive


choice for the sail material as it is RF transparent; however, it is not certain whether it can survive exposure

to atomic oxygen on-orbit [43], [44]. This leaves aluminized Mylar or Kapton as the favourable baseline sail

materials due to its low cost and availability.

However, the aluminum coating on the Mylar and Kapton film is susceptible to humidity in the lab

and at launch sites, causing it potentially to flake off and leave the film exposed once it is on-orbit [46].

Mylar can be considered to have a high moisture sensitivity level (MSL) [42]. Bare Mylar film without a

metallic coating covering it, will not survive on-orbit due to high susceptibility to UV and AO, and if the

aluminum leaves the Mylar substrate, the Mylar will degrade. This problem needs to be better understood

in terms of the timescale and extent of aluminum flaking, as it is may be possible to overcome this problem

with preventative measures (such as a sail bake-out prior to integration, and hermetic or near-hermetic

containment in the spacecraft). Kapton does degrade as Mylar does in a space environment. Kapton can

handle almost all the on-orbit environmental threats, except for atomic oxygen. Therefore, aluminizing the

Kapton will prevent it from degrading due to AO and if the aluminum layer chips off, the Kapton has an

better increased probability of surviving than does a Mylar substrate. Therefore, aluminized Kapton stands

as the baseline choice for the sail material on the CanX-7 mission. Once the sail has been space qualified,

and a more reliable sail material is needed for lifetimes around 25 years, the CP1 material can be purchased

as it has been shown to stand up to the outer space environment during several tests [39], [44].

The Miura-Ori was chosen as the sail fold due to its versatility in shape while the drag sail was being

designed. However, during deployment tests it was found that the fold, with an angle greater than 0◦,

acted like a spring and the booms could not fully deploy the sail material, seen during several deployment

experiments in Figure 4.21.

(a) Fold Acting as a Spring (b) Not Completely Unfolding (c) Encountering Air Resistence

Figure 4.21: Problems with Miura-Ori Deployment Tests

There were several problems during deployment of two different Miura folds shown in the figures. The

first was the air resistance as the device deployed which was also encountered on subsequent testing and is

discussed further in Section 5.5. The result of air resistance on deployment can be seen in Figure 4.21c. The


second problem encountered was with a larger angle on the secondary fold which caused the sail material

to behave like a spring and prevented the material from fully deploying, seen in Figures 4.21a and 4.21b.

The situation may have been aggravated by the air resistance experienced on deployment, but nonetheless

created a problem independent of the cause. The force required to unfold the material from its incomplete

deployment state may be too large to unfold from just the forces experienced on orbit. This will decrease

the usable drag area, particularly at high altitudes where it will significantly affect the resulting lifetime.

Therefore, a variation of the fold with no angle on the secondary fold was tailored specifically to the most

recent drag sail design at the time this thesis was written.

The final variant of the Miura-Ori is seen below in Figure 4.22 and was folded using a two stage process

to try and remove as much air trapped between the folds in order to take up as little volume as possible

and make the folding process repeatable. The fold was done on an aluminized Mylar material and is held in

place in Figure 4.22b using tape that will not be present in flight versions.

(a) Partially Folded (b) Fully Folded

Figure 4.22: Drag Sail Material Fold

Figure 4.23: First Stage Folding Machine

A folding machine, designed by SFL student Vincent Tarantini, was constructed to try and make the

folding process more precise, as folding the sail material by hand introduced small variations amongst the

final folded sails and was labourious and an inefficient use of time. The folding machine consists of two

separate stages, seen in Figures 4.23 and 4.25. The first stage folded the sail in a z-fold, with a 1 inch height,


which is the height of the new sail material allocated volume, which is a triangular section at the opening of

the module that now has a door included in the design, seen in Figure 4.33. The result of the first stage of

folding can be seen in Figure 4.24, where the material used was an non-aluminized Kapton. The resulting

strip of Kapton is held in place by tape to make it easier to move to the second stage of the folding process.

Figure 4.24: Resulting Fold after First Stage

The second stage of the folding machine, seen in Figure 4.25, involves doing another z-fold to make folds

perpendicular to the initial ones. However, the spacing of the secondary folds is not consistent, instead the

folds are spaced to take advantage of the triangular volume drawn onto the machine and used as a guide.

The fold is then slid into its cartridge, the door protecting the sail material is closed and the cartridge is

loaded in between the booms of the device.

Figure 4.25: Second Stage of Folding Machine


4.3.4 Booms

The booms of the drag sail device hold the sail material in place during the deorbit phase, as well as help

the material deploy from its stowed configuration. From initial prototyping both the carbon steel tape and

a COTS measuring tape were used due to their ideal properties of being able to store tension and smoothly

deploy from a stowed configuration. To determine the proper option for the booms a trade study was done,

shown in Table 4.3.

Table 4.3: Boom Material Trade Study

Trade/Option COTS Double COTS Re-bended COTS Custom Carbon Fibre

Manufactured Composites

Cost per Sail 10 20 n/a n/a n/a

($/Sail)

Cost ($) 10 20 n/a >1500 n/a

Lead Time 0 0 n/a n/a n/a

(weeks)

Mass per 40.46 40.46 80.92 25 very

Unit Length low

(g/m)

Volume per 5.46 5.46 10.92 3.64 n/a

Unit Length

(cm3)

Material Stainless Stainless Stainless Variety Carbon

Selection Steel Steel Steel Fibre

Available

Motor No No No No Yes

Required

Of the several options presented, the COTS measuring tape blades were selected as the best option.

It was low in cost, mass, volume and the available tension energy made the use of a motor to drive the

deployment unnecessary. However, obtaining the proper dimensions and ensuring that any two tape springs

purchased would be the same was a concern. The possible inconsistency in manufacturing and some of the

unknown material properties were decided to be mitigated during testing of the booms and the device. The

other options either presented a problem with one of the trades or were not as suited to be used in the


drag sail device. The double COTS solution would provide increased stiffness but would take up too much

volume. The re-bended COTS and custom manufactured options presented were too expensive to use for

multiple modules and the carbon fibre composite option presented a problem with including another moving

device to power the deployment.

Seen below in Figure 4.26 is a photo of a COTS spring steel tape used in prototyping work. With a

change in the module dimensions, see Section 4.6, a larger tape spring was utilized, which provided increased

stiffness. The increased stiffness allowed for better 1g testing and increased stored tension energy for the

device to deploy.

Figure 4.26: COTS Boom

4.4 Prototyping

All stages of the conceptual designs were tested in order to determine the validity of the idea or to see

whether an idea could be made to work. The first prototypes to be made were for the original circular drag

sail unit concept. A small version of the prototype was first made as it was easier to fold the material and

demonstrate the concept before moving on to a larger scale, seen in Figure 4.27. The prototype was made

using a 12.7 μm uncoated Mylar for the sail material, folded using the Miura fold, with a carbon steel raw

spring tape material used for the frame. The small round prototype was found to work and deploy quite

well. The frame was folded, using the initial concept described in Section 4.2.1, and the stored energy in the

frame able to unfold the sail material.


Figure 4.27: Round Small Prototype

Figure 4.28: Round Large Prototype

Therefore, a larger prototype was made, seen in Figure 4.28. The large prototype experienced problems

not seen in the smaller version. First, although the calculated volume of the frame material could easily


fit in the allotted module volume, it was difficult to get the folded frame to fit in the enclosure. This was

because of the problem of the frame material bend radius, a potential problem brought up during initial

concept exploration. The bend radius of the carbon steel tape spring was too great to allow the entire frame

to easily fit. Therefore, moving to a larger drag area if found necessary would prove impossible. Moreover,

because of the circular frame, it was more difficult to stow and deploy the larger frame. The material stowage

and deployment could not be properly scaled between different prototype sizes and the entire concept was

becoming increasingly difficult to work with.

Therefore, a new design concept using the coiled tape springs with the sail material stowed separately

was explored next. To validate that the stored tension device would be capable of deploying and dragging

out the sail with it, a simple prototype was created. The reel seen in Figure 4.29 is oversized and was

intended to prove that a commercial off-the-shelf (COTS) tape spring would be able to deploy under its own

stored energy after being wrapped around the reel. The drag sail was intended to be stored in a separate

compartment above the reel such that it could be dragged out by the tape spring booms but it would not

get caught in the rotating reel, as seen in Figure 4.30. The sail material as folded is slightly too large to fit

in the required dimensions, however, it could be folded to fit. The sail and the booms are attached in this

prototype using strings.

Figure 4.29: Large Reel Prototype

The sail material was folded using a Miura fold. The prototype was made out of plastic, using a rapid


prototype machine. This provided a low cost prototype that was manufactured quickly and was utilized for

several following prototypes quite successfully. The prototype seen in Figure 4.29 was made with a height of

1 cm, therefore the COTS tape spring used was one of the smallest manufactured. The prototype was able

to deploy and showed that the stored tension concept was sound. However, because of the size of the tape

spring and the larger reel, the booms were not fully capable of deploying an attached sail. Also because of

the size of certain components in the design, like the peg that the reel spun around, and the rapid prototype

material not being as strong as the intended machined aluminum, the prototype broke before more testing

could be completed. Nevertheless, there was a lot of knowledge gained through this prototype which would

be applied to future prototypes and designs.

Figure 4.30: Large Reel Prototype in a Stowed Configuration

The next prototype consisted of an increased drag sail module height to 1.6 cm. This allowed a larger

COTS tape spring to be use, and the reel size was also smaller, therefore, the stored tension energy was

greater. The path of the booms was also changed to allow more space for the sail material, seen in Figure

4.20. However, during deployment experiments, it was seen quite quickly that there was a lot of friction

not allowing the booms to initially deploy. When the booms were routed through the new boom paths the

booms were not able to deploy at all and it was even difficult to manually pull them out. With the use

of Teflon tape and carbon fibre lubricant, the booms were able to deploy from the stowed configuration,

but not according to predicted deployment kinematics, further discussed in Section 4.8.2. It was therefore


determined that routing the booms through specific paths was not the solution and was solved in the next

design iteration.

The following group of photos in Figure 4.31 show the deployment sequence of the updated concept that

proves the stored energy concept can work. During the deployment, one of the tape spring booms buckled, in

the top right photo. However, the stored tension energy in the device was enough to overcome the problem

and the deployment continued. The specified boom paths were not used because it determined that the

tape springs required straight paths to deploy properly. A Miura fold was used for folding the sail material

and the booms were able to deploy the fold as they uncoiled off the spool. From this deployment test, the

concept for the new design of the drag sail device was validated and the specific issues hindering deployment

were identified and addressed in future prototypes.

Figure 4.31: Coiled Boom Concept Validation - Deployment Sequence

4.5 Sail to Boom Attachment

The attachment point between the sail material and the booms and the sail material and the module is the

point through which the material is unfolded and how it stays in place during the deorbit period. The survival

of both these attachments is very important to make sure the drag area is not lost and the deorbit period

is not lengthened as a result. Both attachments are envisioned to be similar, however, one major difference

exists. For the sail to boom attachment, it was determined to be better for testing if the attachment could

be removed so that the sail could be refolded and stowed again and then reconnected prior to deployment.

For the sail to module attachment, it was not necessary to have a removable connection and is more practical


to have a more permanent attachment so that the connection is ensured to not fail when the sail is deployed

and pulls on that location. Figure 4.32 shows the initial concept of using a grommet for the sail to boom

attachments.

A grommet will be installed in the sail material and a thread will be the connection to the reattach-

able clasp. An extra two layers of Kapton material will be added around the grommet so that the stress

experienced at that point, when the sail is pulled in tension, will be reduced and spread over a thicker

area. The Kapton will not be glued on as there are no adhesives that will survive the extreme temperature

range seen on-orbit. The grommet shown in Figure 4.32b will not be the final flight attachment, it is shown

here for visualization purposes. Likewise, the lobster clasp, shown in Figure 4.32c is not the final choice

for the removable attachment. Most likely a more secure and strong attachment will be utilized. There is

no dimension restriction on the inner diameter of the grommet, only on the outer diameter to be less then

1 inch as it will need to be folded with the sail material and fit into the sail material cartridge enclosure.

Both the materials for the grommet and thread will be chosen so they can survive the storage period prior

to deployment and the harsh space environment during the deorbit phase.

(a) Sail Attachment (b) Grommet (c) Lobster Clasp

Figure 4.32: Sail to Boom Initial Attachment Concept

4.6 Change of Device Dimensions

Due the change of probable placement of the devices on a GNB and the freeing up of space on a 3U bus,

the initial dimensions of the sail were changed to accommodate using a COTS solutions. One drag sail

module will be 3 cm in height, but as opposed to using the entire 10 cm× 10 cm area it will use half, cut on

a diagonal that allows two modules to be placed side by side when stacked on a 3U, therefore the original

height allocation of 6 cm for the drag sail modules remains. The new triangular shape is seen in Figure 4.33.

The module presented in Figure 4.33, completed by SFL student, Jesse Hiemstra, was a continuation of the

work presented herein.


Figure 4.33: New Drag Sail Housing (Courtesy of SFL)

4.7 Drag Sail Placement

Determining where to place the drag sail modules on a 3U satellite was done with the intent to minimally

interfere with the other satellite components. During deployment, the drag sails also require a clear deploy-

ment plane so the deployment is not hindered and the drag sail components are not damaged. Below are the

placements of the drag on the CanX-7 satellite and the probable placement on the GNB and NEMO SFL

form factors.

4.7.1 Stacked Configuration

The drag sail modules will be stacked on a 3U bus as shown in Figure 4.34. The modules will be placed

on the opposite face of where the antennas are mounted, to ensure no interference upon deployment. The

new shape of the modules still allows mounting to a 3U bus without any deployment problems due to the

satellite legs, shown in the four corners of the satellite. There is space between each of the modules, however,

deployment tests will be completed to ensure adjacent modules do not interfere with one another during

deployment. This potential problem will be investigated in Section 5.5.1.


Figure 4.34: Drag Sail Placement on a 3U (Courtesy of SFL)

4.7.2 Tiled Configuration

The drag sail modules are envisioned to be tiled on the GNB and NEMO satellite buses, shown in Figures

4.35 and 4.36. The exact placement and integration to these two buses are not yet fully determined. There

may also be a structural issue with placing the drag sail modules on the solar array of the NEMO bus. These

and other vital details will be determined once the drag sail has been qualified in space and is being placed

on future SFL satellites.

Figure 4.35: Drag Sail Placement on a GNB (Courtesy of SFL)


Figure 4.36: Drag Sail Placement on a NEMO (Courtesy of SFL)

4.8 Structural Capabilities

The drag sail is required to survive upwards of a 25 year deorbiting period on-orbit. As well, the satellite

is required to undergo some testing in a 1g environment, see Section 5 for more detail, in addition to be

handled, stowed and transported to the launch site under 1g and potentially higher loading. Therefore, it

is necessary to calculate the drag force that the deployed drag sail will see on-orbit and the deployment

kinematics of the sail during testing to better understand how it functions.

4.8.1 Drag Force

Calculations were done to determine the amount of drag force that would be seen on the drag sails as a

satellite deorbits. The drag force was calculated using [28]:

Fd =1

2ρv2CdAd (4.2)

Where, ρ is the density of the atmosphere (kg/m3), v is the satellite velocity (m/s), Cd is the drag

coefficient and Ad is the drag area (m2). The density was obtained from STK, the drag coefficient was

assumed to be 2.2 and the drag area was assumed to be 2.0m2. The velocity, dependent on the altitude was

determined using [29]:

v =

√μ

h(4.3)


Where μ = 3.98× 1014 m3/s2 is the product of the gravitational constant and the mass of the Earth and

h is the orbit altitude (m). Plotting the obtained drag force data, the graph in Figure 4.37 was obtained

over different altitudes. The graph plots the data obtained for a drag area of 2.0m2, a drag coefficient of 2.2

and the speed of the satellite calculated for the specific altitude using Equation 4.3

0.0000001

0.000001

0.00001

0.0001

0.001

0.01

0.1

1

10

100

900 800 700 600 500 400 300 200 100

Force(N)

Orbit Altitude (km)

Figure 4.37: Drag Force Experienced by the Drag Sail at Different Altitudes

As expected the drag force on the drag sail increases as the altitude decreases and the satellite gets

closer to reentering Earth’s atmosphere. Between 200 km and 300 km the drag force is approximately the

force the modules experience on Earth under their own weight. Therefore, the requirement was placed on

the drag sail device to survive up to an altitude of 300 km (TBC) pending any additional analysis. Below

this altitude, a satellite would reenter with or without a drag device in a matter of days and therefore a

functioning drag sail is not required. Making the requirement that the drag sail survives beyond this point

would be overconstraining the design and impractical. Additional analysis needs to be completed to order

to determine what forces would be acting on the connection points, that would be stressed from the drag

force, and if the connections can survive up to this point.

4.8.2 Deployment Kinematics

Each unit of the drag sail device will be stowed in tension and deployed based on a signal sent from an

SFL ground station. In an effort to better understand the deployment process, knowing how much energy

is stored in the booms when they are coiled and how quickly they deploy is very important.

The following Equations, 4.4 - 4.6 can be obtained from [47], which defines the fundamentals of Storable

Tubular Extendible Member (STEM) mechanics. The following calculates the stress found in the booms


when they are in their coiled configuration:

σx =E

1− ν2t

d0

(d0D

− ν

)(4.4)

Where E is the Young’s Modulus (Pa), ν is the Poisson’s Ratio, t is the boom thickness (m), d0 is the

diameter of the spool on which the booms are wound on (m) and D is the boom curvature diameter (m).

For the drag sail, d0 is assumed to equal D, and therefore the ratio, d0/D, is 1. The calculated stress in the

booms was found to be approximately 0.8GPa. This is quite high given that the yield strength for most

stainless steel materials is lower. However, the material does not yield because the spring tape being used is

a carbon steel heat treated such that it can be used as a spring. The stored energy, calculated from Equation

4.5, is 84Nm and is given by:

V1 =E

1− ν2γπt3 (l − x)

6d0

(1 +

d20D2 +

2νd0D

)(4.5)

Where γ is defined as γ = 1+ απ and α is the angle over which the booms are assumed to overlap, which

can be defined negative in this case where the booms are an open arc. For some STEM antennas once

the coiled boom is deployed, the boom will assume its circular shape usually with some overlap instead of

forming a complete or open circle. This provides extra stiffness, not seen with an open cross section. For

the drag sail booms, an overlapped boom configuration is not used because of the increased volume of the

stowed booms and the attachment of the sail material to the booms that may be damaged when the circular

shape is taken. Using the stored energy obtained from Equation 4.5, the speed of deployment of the booms

can be calculated using Equation 4.6, which is the kinetic energy expression, rewritten to solve for velocity:

Ek =1

2mv2

v =

√2Ek

m(4.6)

Where v is the deployment velocity (m/s), and m is the mass of the booms (kg). The calculated speed of

the boom deployment is approximately 55m/s. This speed suggests that the deployment of the booms is very

quick. From further design iterations using ball bearings surrounding the spool mechanism to ensure minimal

friction, the deployment speeds due reach quite high speeds. The earlier prototypes, therefore, experienced

high friction due to small reels, binding of the tape spring to the housing constraints of the reel and boom

paths and were not capable of deploying properly. However, the method of using stored tension energy

through the use of a COTS tape spring was shown to work, both theoretically and through experiment. It


was therefore important to calculate the expected deployment kinematics to ensure no surprises would be

experienced during on-orbit deployment where friction is lower.

4.9 Thermal Analysis

As the drag sail modules are a new technology to SFL, it is necessary to do a thermal analysis that focuses

specifically on the thermal susceptibilities of the device. The large temperature ranges that the modules

will see will affect how the sail deploys and functions. Of particular concern is the maximum three-year

stowage lifetime requirement that the drag sails must be able to survive as they are meant to deploy at the

end of life of what may be a long mission time. The high temperatures and vacuum environment may cause

the tape spring booms to encounter relaxation and lose some of the stored energy and encounter difficulties

during deployment. Moreover, once deployed, the drag sail materials are required to survive the harsh space

environment which includes extensive thermal cycling as the satellite orbits.

4.9.1 Relaxation Analysis

Owing to the drag sail device being stored in tension and experiencing high temperatures over a long period

of time, there arises the possibility that the booms of the drag sail device might experience relaxation and

lose some of the stored tension that will end up deploying the device. For the CanX-7 mission, this might

not be a cause for concern because the secondary payload, operating prior to the drag sail deployment, will

be operational for only 6 months. It is difficult to obtain information specific to the type of spring steel being

used on CanX-7, especially the behaviour on-orbit. However, a plot was obtained, seen in Figure 4.38 that

shows only a small percentage of relaxation will occur for the expected stress loads experienced on the drag

sail over several different temperatures. This graph suggests that for the short secondary payload operation

on the CanX-7 satellite the drag sails will not experience a significant problem due to relaxation.

However, as a general drag sail device requirement, the modules shall survive in a stowed configuration

for up to 3 years, and this could be a challenge when considering how many hot cycles are experienced during

this time period. The concern is, at the end of a potentially three-year period, the coiled tape spring booms

will no longer have the required tension energy to deploy and not be capable of unfolding the drag sail so

that the deorbit of the satellite can begin. If this were the case, the satellite would stay in orbit for a longer

than expected time.

There are several approaches to take in order to mitigate the possibility that relaxation will become a

problem for stowage lifetimes of upwards of 3 years. The situation could be analyzed through simulations or

through accelerated or real time experiments. The latter was chosen after it was discovered that predicting


the behaviour of the tape springs over an extended period of time was a very difficult task. Therefore,

through the use of experiments it will be determined whether the relaxation of the booms will prove to be

a problem for long stowage periods.

Figure 4.38: Relaxation Curve for Carbon Steel [48]

4.9.2 Deployed Drag Sail

Owing to transitioning between large temperature variations on-orbit, long thin structures can undergo

thermal bending. This phenomenon was experienced on the Hubble Space Telescope (HST) after it was

initially placed in orbit. Owing to the severe bending, as can be seen in Figure 4.39, the booms were

replaced during a servicing mission to the HST. There are some concerns that as the drag sail transitions

from Earth’s shadow into full sunlight, and then from sun to shadow several times over the period of a day,

the booms will bend and put stress on the sail material and the connections to the booms. The bending

problem is magnified due to the low torsional stiffness of the booms considering they consist of an open

cross-section [50]. After an initial investigation into the problem, it was found to be quite challenging to

quickly obtain a solution and more analysis is required to determine if this will be a problem or not.


Figure 4.39: Hubble Space Telescope Booms Bending [49]

The sail material also needed to be chosen such that the extreme temperature ranges would not damage

the material. Kapton can survive the large thermal changes, and the aluminization of the Kapton ensures

that it will not experience thermal issues. Aluminized Kapton is also utilized as multi-layer insulation (MLI)

on different spacecraft and has been proven to survive the large thermal cycles.

Chapter 5

Drag Sail Testing

Proper testing of the drag sail device will ensure a thorough design and successful on-orbit functioning.

Because the drag sail is a deployable device, there is a greater risk of a problem occurring as it needs to

be stowed prior to deployment and go through the launch of the satellite and then survive the harsh space

environment prior to and after deploying. Testing will mitigate the risks that will be seen through the life

of the drag device. This chapter will present the test plan envisioned for the drag sail module, as well as

describe deployment and risk mitigation tests done with several prototype models. The test plan presented

here is keeping with SFL guidelines and draws heritage from the eXoadaptable PyrOless Deployer, or XPOD,

system developed at SFL as the XPOD device also consists of a deployable device that releases the satellite

once the launch vehicle has reached its target orbit.

The following sections will describe what tests will be done on the drag sail device to ensure it will

perform as expected when deployed on-orbit. A qualification and acceptance approach will be used to test

the device units. This approach is favored over a protoflight model as the device is a new technology that

has not been widely documented and analyzed. The models undergoing qualification tests will be tested

beyond the expected limits required to survive to or experienced by the actual device to ensure failure will

not happen; or in the case failure occurs, in order to gauge where the failure happens. These models will

not be used as flight models and will only be used in testing. The flight model testing will be similar to the

qualification model testing with some exceptions and will not be nearly as exhaustive. The acceptance tests

will only be done on units intended for flight to ensure that the device will function like the qualification

models if tested further, but to not over stress the units such that they will perform less than ideal or not

at all when on-orbit.

73

Chapter 5. Drag Sail Testing 74

5.1 Qualification Tests

The qualification tests will be done on several units of the device in parallel in order to collect statistics

about the testing and observe the effects of the build process on performance. Qualification testing is used

to test a device beyond its expected working limits in order to ensure all possible methods of failure are

identified and appropriately solved or mitigated. Each qualification test is different and the number of test

cycles completed at each level will vary according to whether it is practical to test the device to failure.

As four units are needed to make a complete drag device, performing qualification tests on at least four

units is desired. Constructing and performing tests on four units may prove difficult in terms of cost, build

time and testing time. Therefore, constructing one or several engineering models to ensure construction will

be consistent and performing qualification tests on the final engineering model is considered an option once

the design is finalized. However, because there are many tests that need to be done, having more qualification

units will allow some testing, where all four units are not required to be tested, to be done in parallel.

Below is a description of the tests envisioned for qualification testing. The tests are not necessarily

presented in the order in which they will occur.

5.1.1 Standard Deployment Test

A standard deployment test will demonstrate the functionality of the drag sail in a 1g environment. Although

at high altitudes the drag force on the drag sail is a fraction of a newton, seen in Figure 4.37 in Section 4.8.1,

deployment in a 1g environment will determine the weak points in the design, which can be strengthened

and will make the design more robust. The test will consist of deploying the drag sail from its stowed

configuration with the coiled booms and folded material. The device will be deployed onto a horizontal

surface to not add additional loads to the design and may require additional ground support equipment

(GSE) to get a realistic on-orbit deployment. During prototype testing, deployment tests were completed

without electronics incorporated into the design in order to test concepts. However, during qualification

tests, electronics will be completely integrated into the drag device to ensure a realistic deployment is being

observed.

An additional deployment test will include deploying an unsupported sail to demonstrate the device

survivability at a low altitude where the drag force will be equal to a 1g environment found on Earth

(Req:CX7-PAY-11, [41]). Similar results can also be gained through a 1g static test, where the drag sail

is deployed on a horizontal surface and then lifted from the surface to demonstrate that the booms do not

buckle. However, putting the constraint that the sail shall survive either test is impractical. The test will

be done for demonstration purposes where success will give a good confidence in the design, but failure will


not invalidate the design as the expected design loads for a large majority of the drag sail life will be well

below the 1g mark.

5.1.2 ReWind Test

Although a rewind capability is not required as a drag sail function on-orbit, it is necessary for testing

the device. After each deployment, it is necessary to restow the booms and refold the sail material. The

rewind of the booms will include a separate motorized device to ensure the booms are not overstressed by

manually forcing them back into their coiled position. The test will be primarily an observational test done

in conjunction with the deployment tests to see how the rewind of the deployment mechanism is affecting

the deployment, if it is at all. This test will provide insight into the condition of the booms over a long

period of time as damage will be evident through the deployment of the device. If the booms buckle and

can not store the required tension energy prior to deployment, the device may jam or not unfold the sail

material properly.

The rewind function will also be required to not damage the inside mechanisms of the drag device and

any sensors that may be present in the device.

5.1.3 Repeated Deployment Tests

The performance of the drag sail over an exhaustive number of trials will determine the cycle life. Repeated

deployment will exhibit the number of times the drag sail can be deployed before the materials no longer

behave as expected and the drag sail can no longer reliably deploy. The number of deployment tests will be

set to 25×’s (TBC) for each unit. The number was chosen based on the need to observe the behaviour of

the device during deployment tests and to see whether failure will result. If failure does not result from the

repeated tests it will add confidence that the design will not fail on-orbit.

However, it may prove to be difficult to refold the sail material and perform further deployment tests and

be confident that the results are the same. Refolding the material presents a problem as new folds may be

introduced during the refolding process and old folds may not be folded the same way. This inconsistency

may result in different deployment dynamics and the repeated deployment test may not be representative

of a typical deployment. Observing whether refolding the material will cause problems was completed with

deployment experiments and will be described in Section 5.4. This is a difficult issue to test as the refolding

may be different every single time and may warrant the use of an expendable sail that will be folded only

once and deployed only once. However, an expendable sail brings up several problems with unit cost of a sail,

time taken to fold and whether testing a new sail every time is realistic in obtaining confidence in repeated


deployment tests. This also brings into concern whether a newly folded sail is sure to deploy if it has not

been tested before hand, bringing up problems if the expendable sail route is chosen for the flight units.

5.1.4 Vibration Test

Vibration testing is necessary as there are many parts that can shift in the device and hinder deployment.

During satellite delivery and launch, the drag sails will experience loads that may cause sensitive parts to

move. Vibration testing will determine if the device can hold up during the expected loads and make it

safely into orbit.

The drag sail device will need to be vibration tested in each axis direction to give confidence it will

reliably deploy. It remains to be seen if all four units need to be tested in each axis direction as an entire

device or if testing one unit in each direction is sufficient. Ideally each qualification unit will be tested in

each axis in order to collect enough data to be certain the device will survive delivery to the launch site and

launch itself. Deployment tests of each vibration tested unit will be done after the vibration test to ensure

the device is functioning properly.

The vibration testing for the drag sail units will follow a similar plan to the XPOD testing [51]. These

test values come from the loads expected on the launch vehicle and delivery to the launch site. Therefore,

the drag sails shall see the following vibration tests:

− Sine burst: 12.3g

− Random: 12.9g RMS

− Shock: 50g

− Modal low-level sine sweeps: 5− 10Hz: 1.0− 4.0g, 10− 100Hz: 4.0g

Vibration tests will not be done on individual flight units, but as an entire device at a spacecraft level

vibe once integration to the bus has occurred.

5.1.5 Thermal Vacuum Test

This test will be performed to ensure the materials are able to handle the extreme temperature variations

and survive the environment of outer space. The thermal vacuum test will consist of:

− Hot Soak

− Hot Start: Deploying the drag sail in a hot environment

− Cold Soak

− Cold Start: Deploying the drag sail in a cold environment


− Sail Stirs and Cycles: Randomly deploying the sail at different temperatures

− Additional deployment tests

The order of the tests remains to be determined, as is the number of units that will go through the

thermal vacuum test. It also remains to be seen if the entire drag sail will be able to completely deploy

within the SFL vacuum chamber without the use of specialized GSE equipment due to the 2m × 2m cross

sectional area. It also remains to be seen if the vacuum seal will be broken to refold the sail and then test

in a different temperature environment. The logistics of the test have not been completely determined due

to the unique nature of having to restow the drag modules after deployment.

5.1.6 Radiation Test

The electronics used to deploy the drag sail device will undergo radiation testing if there is sufficient need.

As they remain unpowered during the mission lifetime and are only active during sail deployment, this test

may not be necessary. However, it remains to be determined what sensors will be used on the drag sail

device and whether some or all of them would remain active during the mission life. If the sensors were

required to stay on for the duration of the satellite mission, there may be just cause to radiation screen the

electronics to show there is a high probability they would survive during the phase prior to sail deployment.

This test would occur prior to testing of the qualification models and would be done at the board level only

and would not require a complete unit. The electronics would be subjected to 15 krad(si) desirable with no

nonresettable failures at a location capable of testing to these specifications.

5.2 Acceptance Tests

Acceptance tests will be done on flight models. Four units plus one flight spare will undergo acceptance

tests. Acceptance testing will be less extreme than the qualification tests and will consist of slightly different

tests that are done to make sure the flight unit will perform like the qualification unit if subjected to further

testing, but is not meant to over stress the unit.

5.2.1 Standard Deployment Test

A deployment test will be performed on each flight unit to ensure the constructed unit is functioning. Two

to three (TBC) tests will be performed using the sail electronics to drive the deployment. It remains to be

seen if the sail material is refolded after each deployment test or whether the flight sail is packaged and not

tested. Unlike the qualification deployment tests described in Section 5.1.1, the acceptance deployment tests


will not consist of static 1g tests to prove the booms will not buckle, as that would have already been tested

with the qualification units. The acceptance deployment tests may also consist of dedicated GSE equipment

to better approximate an on-orbit deployment for the booms with as little friction as possible.

5.2.2 Rewind Test

Again a rewind mechanism is not a required device function. However, the drag device will be subject to a

spacecraft level vibration once the entire spacecraft has been assembled and tested. It will be required to

deploy the drag devices and ensure at this point that they are in fact functioning. The rewind test will most

likely be done after the first deployment test is performed with the procedure worked out from the testing

done on the qualification models. However, if sufficient confidence can be obtained through the qualification

testing that the units will deploy after a vibration test it may not be required to test the deployment. This

would mean no sail membrane would have to be replaced if the expendable sail membrane route was taken.

Although extensive testing may not be enough to prove the drag sail will always deploy the first time, and

therefore there is an increased risk associated with that route.

5.2.3 Thermal Ambient Test

The drag sail units will undergo thermal ambient tests to ensure the required temperature ranges seen

on-orbit do not present a problem for the flight units. The test will include:

− Hot Soak:

− Hot Start: Deploying the drag sail in a hot environment

− Cold Soak

− Cold Start: Deploying the drag sail in a cold environment

− Sail Stirs and Cycles: Randomly deploying the sail at different temperatures

− Additional deployment tests.

This is similar to the qualifications tests described in Section 5.1.5, however, those thermal cycles were

performed in vacuum and these will not be. Again it remains to be seen how the testing of thermal variations

will proceed due to the fact that after the module is exposed to a temperature extreme, it will be deployed

and then restowed before being taken to another temperature.

5.2.4 Thermal Shock

The drag sail flight units will undergo a thermal shock test to ensure they can handle large temperature

extremes. This is a workmanship test and will be done for electronics at a board level only. Thermal tests


will be done on the flight units and the procedure is covered under Section 5.2.3. The test will ensure the

integrity of the electronics.

5.3 Long-Term Tests

Several long-term tests will be completed on the drag sail modules that have completed qualification testing.

As the qualification units will no longer be used as flight models that will be launched on the CanX-7 satellite

or any future satellite, they are available for further testing. The long-term tests will include observing the

drag sail units over extended periods of time that are not possible during the qualification and acceptance

tests due to the time constraints owing to the quick development and launch schedule intended for the

CanX-7 satellite.

5.3.1 Long Term Storage

This test will check how the drag sail deploys when it has been under storage for a long period of time in a

1g environment. This test can be done with the qualification models once they have undergone the required

testing and are no longer of use. As several qualification models will be on hand for the test, the units can

be stored under different conditions. For example, one unit can be cycled hot, cold and then left at room

temperature for a given time frame, while another unit can be cycled hot and cold more often and placed at

a temperature closer to the average that will be seen on-orbit.

Acknowledging time constraints in designing and testing the device, accelerated aging tests should be

considered for some of the drag sail materials as a stand alone test. For example, the sail material could

be folded using the flight material and placed under storage at high temperatures and how the material

subsequently unfolded would determine if any potential problems would arise. If necessary, ‘safe by design’

approaches also need to be considered if long term storage tests become impractical, possibly by analyzing

data collected from other satellites using similar materials or by performing theoretical calculations as to

how the material is suppose to behave.

5.3.2 Long Term Vacuum Test

This test will observe how the sail material survives a long term storage under vacuum. This test will need

to be shortened in order to be completed, as the required to survive on-orbit storage time for the drag sail

can be up to 3 years [41] and keeping something under vacuum for such a long period of time will prove

difficult. Storage of 2 weeks may be representative of what the sail material will see over the course of several


years. The line of thought is that any self-bonding of the material under vacuum will occur within a short

period of time, from the time the material is first exposed to the vacuum environment.

As this test can potentially be resource extensive, it can be avoided if the materials are proven to be

immune to vacuum welding through analysis or on-orbit observations.

5.4 Drag Sail Deployment Tests

Drag sail deployment tests are done with constructed prototypes to prove a concept will work and to also

work out any issues that are encountered throughout design to show they can be solved. Several additional

prototype deployment tests were outlined in Sections 4.3.3 and 4.4 that were completed to determine sail

folding and concept validity.

To determine whether the expendable sail route was the correct path to follow, as initially discussed

in Section 5.1.3, or whether it is possible to refold a sail such that it consistently preforms as expected in

multiple deployment tests, a refoldability test was done. Figure 5.1 shows a deployment test with a sail that

had been deployed twice previously and refolded.

Figure 5.1: Re-foldability Testing

The deployment test with the refolded sail was successful. However, it was observed that refolding the

sail by hand did not result in a neatly folded sail. The resulting sail fold was more bulky than the initial

fold, most likely due to the trapped air that still remained in the folds. Therefore, the sail did not easily

return to the sail cartridge volume. This problem can be mitigated by using the second stage of the sail


material folding machine described at the end of Section 4.3.3 to fully fold the sail material. The first stage

of the machine will not need to be used as those set of folds done come together easily. The deployment

test in Figure 5.1 experienced increased friction and as a result the sail deployment was smooth and did not

encounter the air resistance problems that will be described in Section 5.5.1 and can be seen in Figure 5.3.

The increased friction may have been a result of the sail material not fitting in the sail cartridge volume due

to the refolding. Although, the increased friction produced a positive result, it remains to be seen if this will

always be the case as the friction experienced can be different with every refold. To mitigate this problem,

a second folding machine is currently being constructed so that a more consistent fold can be expected and

the refolding process can be better controlled to expect the same result with every deployment.

5.5 Risk Reduction Tests

Risk reduction tests are done during the drag sail device design to ensure that potential problems that may

arise during on-orbit operations will not cause a failure that will jeopardize the demonstration mission. The

following will describe several tests that have been done to close anticipated problems that might occur.

5.5.1 Boom Exclusion Zone

On the CanX-7 satellite the drag sail modules will be deployed one at a time to determine the deployment

dynamics via sensors and to determine the post sail deployment attitude to verify models, described in

Section 3.2. However, future uses of the drag sail modules will not require large time intervals between

drag sail deployments. A potential concern arose over the possibility that if two modules were deployed

simultaneously, the booms may collide and damage may be done to the sails. Several tests were done to see

if this would be an actual problem during deployment. An exclusion zone test setup was made by setting

up tape for one module and observing the deployment, seen in Figure 5.2. The marked up region has the

dimensions of the fully deployed drag sail, as was previously determined in Figure 4.11.

The booms remained in the marked out region during the test. The marked out area was the actual

drag area of 1m2. In Figure 5.2 it appears that the sail height is too tall; however, this is partly due to

the booms being oversized and the resulting slack that was put into the sail to boom attachment used in

the test. Also contributing to the final shape of the sail was due to the sail material acting as a spring and

pulling the booms inward. This brings up the issue of not being able to utilize the entire drag area if the

sail is not completely unfolded during deployment and whether perturbations experienced on-orbit will be

sufficient to keep the sail unfolded. However, for the purposes of the boom exclusion testing, there was no

concern of one sail hitting another during deployment due to the presence of the sail material.


Figure 5.2: Boom Excursion Deployment Test

Figure 5.3: Example of Incomplete Sail Material Deployment

Another potential problem occurred during the boom excursion testing, seen in Figure 5.3. The booms

were unable to provide the force to completely unfurl the sail material when the material got caught in itself

during deployment. The folds made on the sail material were not clean and some folds ended up having

smaller folds around the intersection of a fold crossover, making it more difficult to properly unfold. While


the force needed to separate the booms and unfold the material was very little, it might not be experienced

on-orbit. However, it was observed upon a slower deployment, during the refoldability tests, Figure 5.1,

that the sail did not get caught. The problem was concluded to be attributed to the sail encountering air

resistance during the deployment because the speed of the boom deployment was high. It remains to be

seen whether this will be a problem as the sail will not encounter any air resistance during an on-orbit

deployment, but this problem may reduce confidence in the ability of the drag sail to properly deploy the

sail material with 1g testing in a nonvacuum setting. If it does become a significant issue, friction may need

to be introduced into the boom deployment to slow it down, while ensuring boom deployment will not be

hindered and there is high confidence that deployment will be successful on orbit.

The refoldability tests, seen in Figure 5.1, also demonstrated the natural tendency of the sail to pull the

booms inwards, and as a result, the booms did not spread beyond the allotted space marked up for one

module. To ensure that the booms would not interfere with one another without the sail, another test was

performed with no sail material, seen in Figure 5.4.

(a) Right Boom (b) Left Boom

Figure 5.4: Boom Excursion Deployment Test, No Sail Material

Although the right boom is outside the exclusion zone, Figure 5.4a, this occurs because the sail material

is not attached. The booms can be made to deploy within the marked out area by slightly adjusting the

angle at which they deploy from the module. Therefore, although this test showed that one of the booms was

outside the exclusion zone, it does not cause a significant problem with the design concept as the problem

can be easily mitigated. The boom exclusion zone testing was done by marking up the area the booms were

not to deploy into because there were not four modules to test side by side. Once four modules are being


tested they will be placed in a flight configuration, either stacked or tiled as presented in Section 4.7, to

ensure the modules do not interfere with one another during deployment.

5.5.2 Tear Propagation

Tear propagation tests were done to see whether the stresses experienced by the sail at 300 km (the required

survivability altitude) or at the start of the deorbit would cause a tear, already present in the sail, to

propagate. The tearing could be caused by micro-meteoroid impact, existing debris collision or during

deployment if the sail were to catch while unfolding. The particular area of concern would be the sail to

boom and the sail to drag sail module attachments as the width of the sail at this location is significantly

less. And a tear at this location would result in losing almost a 1m2 drag area and would significantly affect

the expected deorbit lifetime.

Two separate tests were done to determine how a tear would effect the drag sail. The first consisted of

approximating the worst case drag force that would be experienced on-orbit and trying to match that with

fatigue data obtained for the sail material and finding static loads that could be applied to the material to

see if it failed. The second approach was to take twice the drag force and apply it to the sail material by

cycling to this load from close to zero load over a number of predetermined cycles. Both approaches are

described below.

Figure 5.5: Stress versus Fatigue Life for Kapton [52]


The first approach involved finding an SN curve, a graph of cyclic stress (S) against the logarithmic scale

of cycles to failure (N), for the sail material, seen in Figure 5.5. The endurance limit indicated for this graph

was obtained from the graph found in Figure 5.6.

The expected worst case force acting on the drag sail is approximated to be 0.026N at an altitude of

300 km [26]. This drag force corresponds to the force that pulls on the sail material away from the module

attachment and acts as a centrifugal force, pulling the sail material when the satellite rotates. This was

determined to be larger than the drag force acting perpendicular to the sail material, which was approximated

as a magnitude of less then 0.0026N, see Section 4.8.1.

Figure 5.6: Kapton Endurance Limit [52]

Once the worst expected drag force was obtained, the SN curve in Figure 5.5 was used to find a stress

level at which to test the sail material. The stress ratio [52] used in the figure can be defined as:

Stress Ratio =Smax − Smin

Sendurance(5.1)

Where, Smax is the expected maximum stress on the sail, Smin is the minimum stress and Sendurance is

the value obtained from Figure 5.6.

In the case of the drag sail, the expected minimum stress will be approximately zero as the sail is expected

to not be in full tension for a 100% of one orbit. Calculating the stress ratio puts it well under 1 and therefore

no data for the cycle life exists in the SN curve. Therefore, it makes it difficult to calculate the stress that

should be placed on a test coupon in order to show that any tears will not propagate to see if there will


in fact be a problem on-orbit. Therefore, in order to perform a test, a stress level was chosen close to the

endurance limit as a high end conservative estimate. Figures 5.7 and 5.8 show the results from several static

tests. A mass of 1 kg applied statically to a test coupon of Kapton material was determined from the graph

to be an upper limit to the stress that would be experienced on-orbit. The material used during this test, due

to availability at the time, was a 25μm Kapton, twice as thick as the anticipated flight material, therefore,

the mass was doubled to 2 kg to get representative test results.

(a) Test Coupon (b) Tear Origin (c) Tear that did not Propagate

Figure 5.7: Tear Propagation: Centre and Edge Tear Oscillations

The test coupon in Figure 5.7 was subjected to two tears placed in the centre of the material. The

test coupon was able to survive the 2 kg load when applied statically and when oscillations were induced.

However, of particular concern are edge tears as the sail to boom and sail to module attachments are the

most vulnerable parts of the drag sail when operational and will result in significant drag area lost if severed.

Therefore, an edge tear was introduced into the test coupon in Figure 5.7 to observe what would happen

and if the tear would propagate. The test coupon was able to survive a static application of the mass, but

however failed when oscillations were induced.

(a) Test Coupon (b) Tear Origin

Figure 5.8: Tear Propagation: Edge Tear Oscillations

Figure 5.8 also shows a test coupon failing with only an edge tear subjected to several oscillations with

the 2 kg mass. If the test coupons had been able to survive the conservative loads that were applied, it


would have given high confidence that on-orbit, any tears that were encountered would not pose a problem.

However, as the test coupons failed, a further investigation needed to be completed to determine a more

appropriate test that would better approximate the stress that tears would cause in the sail material. Since

an edge tear possesses a sharp end, the stress experienced at the end is almost infinite, therefore, it is difficult

to approximate a method to test the sail material, and being too conservative with the mass estimates used

to test the material is not representative. In order to get a better feeling for the problem, a test setup was

envisioned to cycle a representative flight material with several tears. In this way the actual worst case load

seen on-orbit can be applied to the test coupon directly.

Figure 5.9: Tear Propagation Test Coupon

Figure 5.10: Tear Propagation Test Setup

A test coupon was made for the cyclic tear propagation experiment. The close up of the sail to boom


attachment in Figure 5.9, on the right, is reinforced with extra material to distribute the stress over a larger

area. The attachment is still the same concept explored and described in Section 4.5. The test setup consists

of the test coupon fixed on one end and attached to a spring that is connected to a motor on the other that

cycles the coupon to twice the expected worst case load, shown in Figures 5.10 and 5.11.

The test was done on several different test coupons of Kapton film with added tears to see if an entire

lifetime cycle of 25 years would cause the tear to propagate. The number of cycles was determined by

assuming that the worst case load on the film would be experienced twice in one orbit as the sail and

satellite rotate. Therefore, if one orbit is assumed to be 1.5 hours, then there will be approximately 2 922 000

cycles experienced in a 25 year lifetime, if one were to apply a safety factor of 10 to the calculations to

account for the fact that this is a fatigue test which requires built-in conservatism. Therefore, the motor

used was set to run at 50 cycles per second, so that the cycle test would be completed in 16 hours.

Figure 5.11: Tear Propagation Test Setup - Connection to Motor

To determine how fast the motor was going, a Hall-effect sensor was placed close to the motor, seen in

Figure 5.12. Two tiny magnets were placed on the shaft of the motor and the Hall-effect sensor was placed

close to the shaft and measured the rotation speed of the motor which could be controlled by varying the

voltage supplied to the motor. The hall effect sensor needed to be used as the motor was purchased cheaply

from a supply store and did not come with set specifications. The sensor was monitored throughout the tests

to determine if the motor is rotating at a constant speed. Several variations in speed were noticed and the


test was allowed to continue a little longer than 16 hours to ensure a 25 year lifetime was being observed.

The first result from the cycle test with the test coupon used in Figure 5.9 was positive. A small tear was

made in the centre of the test coupon, through the load path, and the coupon was cycled for just over 16

hours. The coupon survived the test and the tear did not visibly propagate. From this test, it was desired

to precisely quantify the size of the tear placed in the coupon to be able to draw appropriate conclusions

from the test. After determining the size of the tear to place in the coupon, the tear will be examined and

photographed under a microscope to observe and measure the cut. This will allow any small changes or

propagations that occur during the cycle testing to be spotted and analyzed. Further testing will determine

whether tear propagation will be a significant problem for the sail material and will be mitigated through

the use of ripstop if necessary.

Figure 5.12: Hall Effect Sensor

Chapter 6

Conclusion

Showing that a small satellite will deorbit within a 25 year lifetime is soon being implemented into law. Not

being able to prove a nano- or microsatellite is capable of deorbiting in this time frame will result in delays

and threatens the frequency with which future small satellites will be launched. These satellites, designed

and built under quick schedules and at low cost, are vital for fast-paced demonstrations on-orbit and getting

academic based research into space. The Space Flight Laboratory (SFL) builds nano- and microsatellites

and will soon require a deorbiting device to ensure no delays in future missions.

In an effort to reduce the growing space debris problem, developing a deorbiting device specifically tailored

for the small volume and low mass budgets of small satellites is a worthwhile endeavour. Several recent on-

orbit events have proved that space will only get more crowded unless a method is found to mitigate the

situation. Typical deorbiting methods for larger satellites, like using rockets, are impractical for use on

small satellites. The Space Flight Laboratory is developing the CanX-7 satellite with the primary mission

of demonstrating a deorbiting drag sail technology.

The lifetime analysis shows that a 4.0 m2 drag sail with an effective area of 2.0 m2 will deorbit the CanX-

7 satellite well within the required IADC guidelines of less than 25 years. The lifetime analysis consisted

of looking at different parameters that affect how long a satellite will spend deorbiting. These parameters

included the drag area, the drag coefficient, ballistic coefficient, solar activity predictions, atmospheric density

model, effective area and the launch year of the satellite. Modeling the satellite properties and varying certain

parameters all produce errors. However, measuring the error is also difficult because the uncertainty is not

quantifiable. Therefore, each parameter is varied for a certain group of satellite properties and evaluated

based on how it effects the deorbiting lifetime and a certain error is then given. In this way, an approximate

error is associated with the lifetime calculations and the drag sail is oversized accordingly.

A drag sail was chosen as the most suitable device for deorbiting SFL spacecraft. After several design

90

Chapter 6. Conclusion 91

iterations, it was decided that a noncentralized device would function through the use of stored tension

energy. The drag sail will consist of two booms coiled around a spool, held in place from deploying, and

once a signal is received to deploy, the booms pull and unfold a sail material that interacts with the Earth’s

atmosphere and causes the host satellite to deorbit. Coiling the tape spring booms as opposed to folding was

determined from prototyping and testing. The sail material was chosen to be 12.7 μm aluminized Kapton

film due to its survivability in space, cost and availability.

Testing of the drag sail is done in order to mitigate any potential risks as this is a new technology for SFL.

Testing of the drag sail spans everything from deployment tests, folding and refolding of the sail material,

to thermal cycling the device to ensure it can survive the harsh space environment. The test plan for the

drag sail will be a qualification and acceptance approach, to ensure the device concept is thoroughly tested

prior to building flight models. There still remain several tests that need to be done before the building and

testing of the qualification models. These tests will include ensuring the sail to boom attachment concept

is valid and prototyping the attachment with flight representative materials. As well, continuing to explore

test methods for the tear propagation tests will be done in addition to any testing required to mitigate any

operational risks.

Once the drag sail device has been validated to work on-orbit through the CanX-7 satellite and the device

has flight heritage, it can be further sized to integrate on larger SFL satellites and be used to ensure future

small satellites missions can continue without delay.

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