the odyssey project team members - purdue university · aae450 - senior spacecraft design – fall...

AAE450 - Senior Spacecraft Design – Fall 2006 The Odyssey Project

Final Report

The Odyssey Project Team Members ConOps

• Matt Harvey • David Helderman • Ashley Ruic

Cost & Scheduling • Matt Harvey

Launch Vehicle • Nick Sochinski

Orbital Selection • Stephanie White

Payload – Optics, Data Handling, and Communication • Nick Andrews • Craig Bittner • Matt Dennis • Elisabeth Hanssens • Ashley Ruic

Power Systems • Norman Herbertz

Propulsion & ACS • Jon Fromm • David Helderman • Stephanie White

Structures

• Hadi Ali • April Miller • Aaron Schinder

Systems Engineering • David Helderman • Ashley Ruic

Thermal Controls

• Chris Murphy


Final Report i

i. - Table of Content by Topic i. - Table of Content by Topic.......................................................................................................... i ii. - Table of Content by Author..................................................................................................... vi iii. - Acronyms ............................................................................................................................... xi iv. - Preface by the Professor ........................................................................................................ xii 1.0 – Background: Future Needs .................................................................................................... 1

1.1 – New System Requirements ................................................................................................ 1 1.2 – Fractionated Systems ......................................................................................................... 1 1.3 – Modular Systems ............................................................................................................... 2 1.4 – Mission Opportunity Statement ......................................................................................... 3 1.5 – References.......................................................................................................................... 6

2.0 – Design Approach and Requirements ..................................................................................... 7 2.1 – Space System Design......................................................................................................... 7 2.2 – Determining requirements and Design Drivers ................................................................. 8

2.2.1 – Primary System Requirements.................................................................................... 8 2.2.2 – Secondary System Requirements................................................................................ 9 2.2.3 – Tertiary System Requirements.................................................................................... 9

2.3 – Requirements ................................................................................................................... 10 2.3.1 – Explanation of Traceability....................................................................................... 10 2.3.2 – Tier Requirements..................................................................................................... 10

2.4 – References........................................................................................................................ 14 3.0 – Design Concept.................................................................................................................... 15

3.1 – Brainstorming Design Concepts ...................................................................................... 15 3.2 – Concept Selection ............................................................................................................ 16 3.3 – Final Concept Description ............................................................................................... 18 3.4 – References........................................................................................................................ 20

4.0 – Payload – Optics .................................................................................................................. 21 4.1 – Scope and Purpose ........................................................................................................... 22 4.2 – Tasks, Functions, Requirements and Design Methodology............................................. 22 4.3 – Design Choices and Drivers............................................................................................. 22 4.4 – Design Evolution ............................................................................................................. 24

4.4.1 – Earth Viewing Telescope.......................................................................................... 26 4.4.2 – Space Viewing .......................................................................................................... 35

4.5 – Cost Estimation................................................................................................................ 42 4.6 – Summary – Lessons learned and future systems ............................................................. 43 4.7 – References........................................................................................................................ 45

5.0 – Orbit Selection ..................................................................................................................... 47 5.1 – Scope and Purpose ........................................................................................................... 47 5.2 – Tasks, Functions, Requirements and Design Methodology............................................. 47 5.3 – Orbit Selection - Problems and Issues ............................................................................. 48

5.3.1 – Earth Imaging............................................................................................................ 48 5.3.2 – Space Imaging........................................................................................................... 48 5.3.3 – Trade-off: Simple Plane Change Maneuvers............................................................ 49

5.4 – System Orbit Description................................................................................................. 50


Final Report ii

5.4.1 – Sun-synchronous Orbit (SS-O) ................................................................................. 50 5.4.2 – Orbit Characteristics ................................................................................................. 51 5.4.3 – ΔV for Orbit Insertion............................................................................................... 53 5.4.4 – Launch Window........................................................................................................ 53

5.5 – Summary – Lessons Learned and Outlook for the Future ............................................... 53 5.6 – References........................................................................................................................ 54

6.0 – Communications and Data Handling ................................................................................... 55 6.1 – Scope and Purpose ........................................................................................................... 55 6.2 – Tasks, Functions, Requirements and Design Methodology............................................. 55 6.3 – Choices............................................................................................................................. 57 6.4 – Design Evolution ............................................................................................................. 58 6.5 – Telecommunication Design ............................................................................................. 58

6.5.1 – Design Approach ...................................................................................................... 58 6.5.2 – Requirements – Trade Studies .................................................................................. 59 6.5.3 – Final Concept – Component Descriptions ................................................................ 59

6.6 – Data Handling Design...................................................................................................... 61 6.6.1 – Design Approach ...................................................................................................... 61 6.6.2 – Final Concept – Components and Specifications ..................................................... 61

6.7 – Material Cost Estimation ................................................................................................. 62 6.8 – Summary – Lessons Learned and Outlook for Future ..................................................... 62 6.9 – References........................................................................................................................ 64

7.0 – Structures and Bus Design ................................................................................................... 65 7.1 – Scope and Purpose ........................................................................................................... 65 7.2 – Tasks, Functions, Requirements and Design Methodology............................................. 65 7.3 – Design Choices and Drivers............................................................................................. 66 7.4 – Design Evolution ............................................................................................................. 67 7.5 – Segment Configuration .................................................................................................... 67

7.5.1 – Design Choices and Decisions.................................................................................. 67 7.5.2 – Core Design – Approach, Methodology and Evolution............................................ 67 7.5.3 – Module Design.......................................................................................................... 73

7.6 – Structural Design of the Propulsion Segment.................................................................. 74 7.6.1 – Configuration Design................................................................................................ 74 7.6.2 – Structural Design ...................................................................................................... 75

7.7 – Module Docking Design .................................................................................................. 77 7.7.1 – Scope, Requirements, Choices, Design Methodology and Approach ...................... 77 7.7.2 – Concept Selection ..................................................................................................... 78 7.7.3 – Discussion of Design Choices .................................................................................. 79 7.7.4 – Module Docking ....................................................................................................... 80 7.7.5 – Segment Docking Concept ....................................................................................... 82

7.8– Cost Estimation................................................................................................................. 83 7.9 – Summary and Lessons Learned ....................................................................................... 84 7.10 – References...................................................................................................................... 85

8.0 - Attitude Control System ....................................................................................................... 86 8.1 - Scope and Purpose............................................................................................................ 86 8.2 - Tasks, Functions, Requirements and Design Methodology ............................................. 86 8.3 – Choices - Decision Justification and Design Drivers ...................................................... 87


Final Report iii

8.3.1 – Disturbance Torques ................................................................................................. 87 8.3.2 – Center of Mass .......................................................................................................... 87 8.3.3 – Three-axis Stabilized System.................................................................................... 88 8.4.1 – Rate-sensing Gyroscopes .......................................................................................... 89 8.4.2 - Reaction Wheels ........................................................................................................ 90 8.4.3 - Fine Guidance Sensors .............................................................................................. 90

8.5 – Mass and Power Requirements and Discussion of Cost .................................................. 91 8.6 – Docking Process............................................................................................................... 92

8.6.1 – Locate Module .......................................................................................................... 92 8.6.2 – Direct Module ........................................................................................................... 92 8.6.3 – Dock Module ............................................................................................................ 93

8.7 - ΔV Budget ........................................................................................................................ 93 8.8 - Summary – Lessons Learned and Future Efforts ............................................................. 94 8.9 - References ........................................................................................................................ 95

9.0 – Launch Vehicle System ....................................................................................................... 96 9.1 Scope and Purpose .............................................................................................................. 96 9.2 Tasks, Functions, Requirements, and Design Methodology............................................... 96 9.3 Choices - Trade Studies and Decision Justification............................................................ 99 9.4 Design Evolution .............................................................................................................. 101 9.5 Launch Site ....................................................................................................................... 101 9.6 Primary Launch Vehicles and Environment ..................................................................... 102 9.7 Backup Launch Vehicle.................................................................................................... 102 9.8 Cost Estimate .................................................................................................................... 103 9.9 Summary – Lessons Learned and Future Efforts.............................................................. 103 9.10 References....................................................................................................................... 104

10.0 – Propulsion System ........................................................................................................... 105 10.1 - Scope and Purpose........................................................................................................ 105 10.2 - Tasks, Functions, Requirements, and Design Methodology ........................................ 106 10.3 - Choices - Trade Studies, Decision Justification, and Design Drivers .......................... 107 10.4 - Design Evolution .......................................................................................................... 109 10.5 - Final Propulsion Concept Overview............................................................................. 110 10.6 – Subsystem Design, Specification, and Operation ........................................................ 111

10.6.1 – Altitude Maintenance Subsystem ......................................................................... 111 10.6.2 – Attitude Control Thruster Subsystem ................................................................... 112 10.6.3 – Fuel Distribution Subsystem................................................................................. 114 10.6.4 – Module Propulsion Subsystem ............................................................................. 116 10.6.5 – Super-module Propulsion Subsystem ................................................................... 118 10.6.6 – Segment Propulsion Subsystem............................................................................ 118

10.7 – Propellant Budget and Refueling Schedule ................................................................. 119 10.8 – System Life Expectancy .............................................................................................. 120 10.9 – Cost Estimation............................................................................................................ 120 10.10 – Summary .................................................................................................................... 121 10.11 – References.................................................................................................................. 122

11.0 – Power System................................................................................................................... 123 11.1 – Scope and Purpose ....................................................................................................... 123 11.2 - Tasks, Functions, Requirements, and Design Methodology ........................................ 123


Final Report iv

11.3 – Choices: Trade Studies and Decision Justification, Design Drivers............................ 124 11.4 – Design Evolution ......................................................................................................... 125 11.5 – Solar Array Design ...................................................................................................... 125 11.6 – Battery Design ............................................................................................................. 126 11.7 – Bus Bar design ............................................................................................................. 127 11.8 – Material Cost................................................................................................................ 127 11.9 – Summary ...................................................................................................................... 127 11.10 – References.................................................................................................................. 129

12.0 – Thermal Control............................................................................................................... 130 12.1 – Scope and Purpose ....................................................................................................... 130 12.2 – Tasks, Functions, Requirements and Design Methodology......................................... 130

12.2.1 Space Environment .................................................................................................. 132 12.3 – Thermal Components................................................................................................... 138 12.4 – Trade Studies and Decision Justification..................................................................... 140 12.5 – System Design ............................................................................................................. 141

12.5.1 – Multilayered Insulation......................................................................................... 141 12.5.2 – Propulsion Segment .............................................................................................. 141 12.5.3 – Optical Segment.................................................................................................... 142

12.6 – Cost Estimation............................................................................................................ 143 12.7 – Summary ...................................................................................................................... 143 12.8 – References.................................................................................................................... 145

13.0 – Experiment Segment........................................................................................................ 147 13.1 – Scope and Purpose ....................................................................................................... 147 13.2 – Power Beaming Demonstrator..................................................................................... 147 13.3 – Directed Energy Weapon............................................................................................. 148 13.4 – Nanosat Docking Station ............................................................................................. 148 13.5 – NEO and ELE object detection.................................................................................... 148 13.6 – Summary ...................................................................................................................... 149 13.7 – References.................................................................................................................... 150

14.0 – Cost and Scheduling ........................................................................................................ 151 14.1 – Scope and Purpose ....................................................................................................... 151 14.2 – Cost Analysis ............................................................................................................... 151

14.2.1 Hubble Comparison ................................................................................................. 153 14.3 – Scheduling.................................................................................................................... 154 14.4 – References.................................................................................................................... 156

15.0 – ConOps ............................................................................................................................ 157 15.1 – Scope and Purpose ....................................................................................................... 157 15.2 – Business Plan ............................................................................................................... 157 15.3 – Ground Systems ........................................................................................................... 157 15.4 – Operations .................................................................................................................... 159

15.4.1 – Space Viewing Telescope ..................................................................................... 159 15.4.2 – Earth Viewing Telescope...................................................................................... 159 15.4.3 – Technology Demonstrator .................................................................................... 160

15.5 – Difference between fractionated systems and single function system......................... 160 15.6 – Summary ...................................................................................................................... 161 15.7 – References.................................................................................................................... 162


Final Report v

16.0 – Conclusion ....................................................................................................................... 163 16.1 – Project Summary.......................................................................................................... 163 16.2 – Lessons Learned........................................................................................................... 164 16.3 – Closing and Remarks ................................................................................................... 165 16.4 – References.................................................................................................................... 166


Final Report vi

ii. - Table of Content by Author Hadi Ali (168)

7.1 – Scope and Purpose ........................................................................................................... 65 7.2 – Tasks, Functions, Requirements and Design Methodology............................................. 65 7.6 – Structural Design of the Propulsion Segment.................................................................. 74 7.8– Cost Estimation................................................................................................................. 83

Nick Andrews (168)

4.4 – Design Evolution ............................................................................................................. 24 4.4.1 – Earth Viewing Telescope.......................................................................................... 26

4.5 – Cost Estimation................................................................................................................ 42 Craig Bittner (168)

6.3 – Choices............................................................................................................................. 57 6.6 – Data Handling Design...................................................................................................... 61 6.7 – Material Cost Estimation ................................................................................................. 62 6.8 – Summary – Lessons Learned and Outlook for Future ..................................................... 62

Matt Dennis (168) 6.0 – Communications and Data Handling ................................................................................... 55

6.1 – Scope and Purpose ........................................................................................................... 55 6.2 – Tasks, Functions, Requirements and Design Methodology............................................. 55 6.3 – Choices............................................................................................................................. 57 6.4 – Design Evolution ............................................................................................................. 58 6.5 – Telecommunication Design ............................................................................................. 58 6.7 – Material Cost Estimation ................................................................................................. 62 6.8 – Summary – Lessons Learned and Outlook for Future ..................................................... 62

Jon Fromm (169) 10.0 – Propulsion System ........................................................................................................... 105

10.1 - Scope and Purpose........................................................................................................ 105 10.2 - Tasks, Functions, Requirements, and Design Methodology ........................................ 106 10.3 - Choices - Trade Studies, Decision Justification, and Design Drivers .......................... 107 10.4 - Design Evolution .......................................................................................................... 109 10.5 - Final Propulsion Concept Overview............................................................................. 110 10.6 – Subsystem Design, Specification, and Operation ........................................................ 111

10.6.1 – Altitude Maintenance Subsystem ......................................................................... 111 10.6.2 – Attitude Control Thruster Subsystem ................................................................... 112 10.6.3 – Fuel Distribution Subsystem................................................................................. 114

10.7 – Propellant Budget and Refueling Schedule ................................................................. 119 10.8 – System Life Expectancy .............................................................................................. 120 10.9 – Cost Estimation............................................................................................................ 120 10.10 – Summary .................................................................................................................... 121


Final Report vii

Elisabeth Hanssens (169) 3.0 – Design Concept.................................................................................................................... 15

3.1 – Brainstorming Design Concepts ...................................................................................... 15 3.2 – Concept Selection ............................................................................................................ 16 3.3 – Final Concept Description ............................................................................................... 18

Matt Harvey (169) 14.0 – Cost and Scheduling ........................................................................................................ 151

14.1 – Scope and Purpose ....................................................................................................... 151 14.2 – Cost Analysis ............................................................................................................... 151 14.3 – Scheduling.................................................................................................................... 154

15.0 – ConOps ............................................................................................................................ 157 15.1 – Scope and Purpose ....................................................................................................... 157 15.2 – Business Plan ............................................................................................................... 157 15.3 – Ground Systems ........................................................................................................... 157 15.4 – Operations .................................................................................................................... 159 15.5 – Difference between fractionated systems and single function system......................... 160 15.6 – Summary ...................................................................................................................... 161

David Helderman (169)

10.6.4 – Module Propulsion Subsystem ............................................................................. 116 10.6.5 – Super-module Propulsion Subsystem ................................................................... 118 10.6.6 – Segment Propulsion Subsystem............................................................................ 118

13.0 – Experiment Segment........................................................................................................ 147 13.1 – Scope and Purpose ....................................................................................................... 147 13.2 – Power Beaming Demonstrator..................................................................................... 147 13.3 – Directed Energy Weapon............................................................................................. 148 13.4 – Nanosat Docking Station ............................................................................................. 148 13.5 – NEO and ELE object detection.................................................................................... 148 13.6 – Summary ...................................................................................................................... 149

15.0 – ConOps ............................................................................................................................ 157 15.1 – Scope and Purpose ....................................................................................................... 157 15.2 – Business Plan ............................................................................................................... 157 15.3 – Ground Systems ........................................................................................................... 157 15.4 – Operations .................................................................................................................... 159 15.5 – Difference between fractionated systems and single function system......................... 160 15.6 – Summary ...................................................................................................................... 161

16.0 – Conclusion ....................................................................................................................... 163 16.1 – Project Summary.......................................................................................................... 163 16.2 – Lessons Learned........................................................................................................... 163 16.3 – Closing and Remarks ................................................................................................... 163


Final Report viii

Norm Herbertz (169) 11.0 – Power System................................................................................................................... 123

11.1 – Scope and Purpose ....................................................................................................... 123 11.2 - Tasks, Functions, Requirements, and Design Methodology ........................................ 123 11.3 – Choices: Trade Studies and Decision Justification, Design Drivers............................ 124 11.4 – Design Evolution ......................................................................................................... 125 11.5 – Solar Array Design ...................................................................................................... 125 11.6 – Battery Design ............................................................................................................. 126 11.7 – Bus Bar design ............................................................................................................. 127 11.8 – Material Cost................................................................................................................ 127 11.9 – Summary ...................................................................................................................... 127

April Miller (170)

7.3 – Design Choices and Drivers............................................................................................. 66 7.4 – Design Evolution ............................................................................................................. 67 7.7 – Module Docking Design .................................................................................................. 77 7.9 – Summary and Lessons Learned ....................................................................................... 84 8.6 – Docking Process............................................................................................................... 92

Chris Murphy (170) 12.0 – Thermal Control............................................................................................................... 130

12.1 – Scope and Purpose ....................................................................................................... 130 12.2 – Tasks, Functions, Requirements and Design Methodology......................................... 130 12.3 – Thermal Components................................................................................................... 138 12.4 – Trade Studies and Decision Justification..................................................................... 140 12.5 – System Design ............................................................................................................. 141 12.6 – Cost Estimation............................................................................................................ 143 12.7 – Summary ...................................................................................................................... 143


Final Report ix

Ashley Ruic (170) 1.0 – Background: Future Needs .................................................................................................... 1

1.1 – New System Requirements ................................................................................................ 1 1.2 – Fractionated Systems ......................................................................................................... 1 1.3 – Modular Systems ............................................................................................................... 2 1.4 – Mission Opportunity Statement ......................................................................................... 3

2.0 – Design Approach and Requirements ..................................................................................... 7 2.1 – Space System Design......................................................................................................... 7 2.2 – Determining requirements and Design Drivers ................................................................. 8 2.3 – Requirements ................................................................................................................... 10

4.0 – Payload – Optics .................................................................................................................. 21 4.1 – Scope and Purpose ........................................................................................................... 22 4.2 – Tasks, Functions, Requirements and Design Methodology............................................. 22 4.3 – Design Choices and Drivers............................................................................................. 22 4.4 – Design Evolution ............................................................................................................. 24

4.4.2 – Space Viewing .......................................................................................................... 35 4.5 – Cost Estimation................................................................................................................ 42 4.6 – Summary – Lessons learned and future systems ............................................................. 43

Aaron Schinder (170)

7.5 – Segment Configuration .................................................................................................... 67 Nick Sochinski (171) 9.0 – Launch Vehicle System ....................................................................................................... 96

9.1 Scope and Purpose .............................................................................................................. 96 9.2 Tasks, Functions, Requirements, and Design Methodology............................................... 96 9.3 Choices - Trade Studies and Decision Justification............................................................ 99 9.4 Design Evolution .............................................................................................................. 101 9.5 Launch Site ....................................................................................................................... 101 9.6 Primary Launch Vehicles and Environment ..................................................................... 102 9.7 Backup Launch Vehicle.................................................................................................... 102 9.8 Cost Estimate .................................................................................................................... 103 9.9 Summary – Lessons Learned and Future Efforts.............................................................. 103


Final Report x

Stephanie White (171) 5.0 – Orbit Selection ..................................................................................................................... 47

5.1 – Scope and Purpose ........................................................................................................... 47 5.2 – Tasks, Functions, Requirements and Design Methodology............................................. 47 5.3 – Orbit Selection - Problems and Issues ............................................................................. 48 5.4 – System Orbit Description................................................................................................. 50 5.5 – Summary – Lessons Learned and Outlook for the Future ............................................... 53

8.0 - Attitude Control System ....................................................................................................... 86 8.1 - Scope and Purpose............................................................................................................ 86 8.2 - Tasks, Functions, Requirements and Design Methodology ............................................. 86 8.3 – Choices - Decision Justification and Design Drivers ...................................................... 87 8.5 – Mass and Power Requirements and Discussion of Cost .................................................. 91 8.7 - ΔV Budget ........................................................................................................................ 93 8.8 - Summary – Lessons Learned and Future Efforts ............................................................. 94


Final Report xi

iii. - Acronyms ACS - Attitude control system CCD - Charge coupled device CVZ - Continuous viewing zone DAS - Demand access system EELV - Evolved expendable launch vehicle ELE - Extinction level event EVC - Earth viewing camera FEA - Finite element analysis FGS - Fine guidance sensor FOV - Field of view GA - Genetic algorithm HST - Hubble space telescope IR - Infrared radiation JWST - James Webb space telescope LADAR - Laser detection and ranging LEO - Low earth orbit LVS - Launch vehicle system MLI - Multi layered insulation MLT - Mean local time of ascending node MPS - Module propulsion system NASA - National Aeronautics and Space Administration P&ID - Plumbing and instrumentation diagram PCS - Pointing control system (used on HST) RAAN - Right ascension of ascending node RFP - Request for proposal RSU - Rate sensing unit RWA - Reaction wheel assembly SLC-3W- Space launch complex 3 west SOG - Satellite operations group SS-O - Sun-synchronous orbit STK - Satellite tool kit SVM - Space viewing mirror TCS - Thermal control systems TDRS - Tracking and data relay satellite TDRSS - Tracking and data relay satellite system VAFB - Vandenberg air force base WLGS - West Lafayette ground station WSGC - White sands ground complex


Final Report xii

iv. - Preface by the Professor

Aspirations, background and lessons learned My academic advisor at Stanford wrote in the preface to one of his textbooks, “The preface is the part of a book which is written last, placed first and read least.” It is the only place to explain what the instructor is up to and, more importantly, to acknowledge the supporting cast of people in the AAE School who actively participated in this project. As a result (and true to my own form), this preface is lengthy. The supporting cast for AAE450 is a short but important list. It begins with Geoff Wawrzyniak, the Teaching Assistant for the course. Geoff is young, enthusiastic and skilled, having worked for a time at JPL before returning to school to pursue a Ph.D. degree. His special responsibility is to help (and comfort) students, grade homework and reports (sometimes given to him at the last minute) and to advise the professor (even when the professor doesn’t know that he needs advice). Geoff displayed talent to the degree that he became a partner in the enterprise. Special recognition is also given to Dr. David Filmer who, after his retirement several years ago from Purdue’s EE department, came to AAE to assist in our efforts to develop a first-class astronautics program. Dave is an expert on several critical technical subjects, most importantly space communications and assisted the students in this critical design area. Also deserving special mention is Professor William Anderson. Bill didn’t know how important he was to the student effort. His propulsion design course, taken by several of the design team members, is of such high quality that critical design efforts such as formulating system design requirements needed little explanation to several students engaged in this important part of the design effort. The propulsion effort for this class is of particularly high quality and I credit this success to the strong foundation laid by Bill Anderson. Bill, together with Professor Ivana Hrbud, attended the class final presentation and, by their presence, provided special recognition to our students. The design team performed very well. As always there are a few students who shoulder the extra burden of management and data handling. In particular, this semester there were two students who were chosen by their peers in end-of-class surveys as deserving of special mention. These two students, David Heldeman and Ashley Ruic were co-team leaders. Their team members singled them out for special mention with comments such as “Tremendous respect from peers” and “absolutely amazing energy.” Now you can go to the design report, or you may also want to read about what the design class goals and methods were. Like any effort of this kind there were successes and failures on the part of the professor. We will cover these in what follows. Special considerations – the purpose of a design course Because design courses are unlike academic courses with focused mono-disciplinary objectives, multi-disciplinary design courses are challenge for the professor and the students and the source of misunderstanding by students and faculty. Questions abound. Is design just a cookbook


Final Report xiii

course with a project to be approached like a super homework problem—find the formulas and draw some pictures? What type of design problem is appropriate for Aerospace engineering students—does it matter what the topic is, as long as there is an open ended problem? By the way, what is an “open-ended” problem? Many strong opinions are available, some by faculty who have taught the course, some who have not. Teaching style and educational objectives are strong functions of educational background and professional experience. For the past three decades I have had the good fortune to be an active participant in advanced aerospace development programs, both as a consultant, a member of national advisory boards and as a government program developer and manager. I have used this experience to develop an approach to design that I believe is effective and that has enhanced the opportunities for hundreds of Purdue students that I have taught, although not all students would agree. The central purpose of AAE450 is to use challenging engineering projects to enable student teams to demonstrate that they are capable of executing complex aerospace design tasks and achieving professional results. These objectives require three important elements of contemporary design:

1. Application of formal design methods 2. Management of projects with hard deadlines 3. Effective teamwork and group dynamics

My own overarching goals are:

• To provide a design experience that begins as close to embryonic conception as possible, yet carries the process to a sufficient stage where an innovative, feasible concept is defined to the extent that it can be evaluated for viability. This feasibility assessment the important considerations of cost and manufacturability.

• To provide a challenging environment that moves students outside their comfort levels to work in teams and to acquire new skills and new knowledge beneficial to their future careers. This includes writing reports, as well as preparing and delivering clear, concise presentations during which they are challenged to support conclusions

Because the semester at Purdue lasts only fourteen weeks we are presented with a challenge to achieve these goals. We need to consider key questions such as:

• How does one structure a design class to provide education and skills? • Where does the class begin? • Where does it end?

To provide the answers to these questions, we need to consider the product life cycle phases, shown in Figure 1 on the next page. Designing a design course activity requires trade-offs, just like an engineering design problem. A project embracing the entire product life cycle is not easily covered in one or even two


Final Report xiv

semesters unless it involves a focused product with strong support teams. (This is the case with solar racing teams for instance.) To move as far as the “system implementation” phase with a small design team, limited resources, and limited experience, we need to reduce the level of sophistication of the design problem itself, enter at a level beyond project conception, or reduce technical sophistication. A design class involving a sophisticated satellite system cannot penetrate far into the product life cycle, although students can be challenged to consider all phases of the cycle. Entrepreneurial companies and the government agencies “begin at the beginning” with a “vision statement” or “capabilities statement” consisting of only a few words from a customer “…landing a man on the moon and returning him safely to the Earth.”1 This is sometimes called the “mission opportunity statement” with important details left vague and to be filled in by the design team. This type of challenge is fun—if you do it correctly you will stay in business and prosper.

1 President John F. Kennedy, Urgent National Needs, a speech to a joint session of Congress, May 25, 1961. The words in this speech delivered over 40 years ago still resonate and inspire. Kennedy used phrases like “more money alone will not do the job” and “If we were only to go halfway, or reduce our sights in the face of difficulty, it would be better not to go at all.” These are applicable to almost every new effort with the capability to produce great results.


Final Report xv

Figure 1 - System Life Cycle Phases2

The first phase of a design effort not only includes describing the product need, but also includes preparing a set of requirements at different levels, developing and selecting concepts and then producing a conceptual design, complete with drawing, technical details and operational plans. This effort consumes only about 8% of the total development effort, but locks in about 70% of the product life cycle cost.3 Undergraduate students are well qualified to participate in this kind of activity. More importantly, students learn how projects originate and are challenged to become innovators. Students are also challenged to work outside their own comfort zones and to 2 Systems Engineering Handbook, International Council on Systems Engineering, SE Handbook Working Group, INCOSE-TP-2003-16-02, Version 2a, 2004, p. 26. 3 A Russian engineer, Genrich Altshuller screened over 200,000 patents looking for inventive problems and how they were solved. He found that of the 200,000 patents examined, only 40,000 met his criteria for “inventive” solutions; the rest were considered to be straightforward improvements of other ideas. Altshuller defined an inventive problem as one in which the solution causes another problem to appear. As a result of his studies, he claimed that over 90% of the problems engineers faced had been solved somewhere before. It is the other 10% that present problems and create opportunities.


Final Report xvi

realize that a B.S. degree is a launching point, not a final destination. This goal can produce strong student dissatisfaction and impatience to “start the design.” Teams involved in Phase 1 conceptual efforts are usually small. I believe that five to six creative, skilled members are ideal. However, in the university environment, class technical composition in terms of major areas and minor areas represented is uncontrolled. It is not unusual to find, as we did this semester, that there are not enough people with critical skills to staff more than one team. As a result, the single AAE450 team of fifteen members was larger that I would have liked because only one concept was pursued. Moreover, there was no inter-team class competition. Competition is an effective way of encouraging teams to avoid complacency. Systems engineering processes Process is important to any successful design effort. While students have very limited technical skills, have very little practical engineering experience and absolutely no managerial experience, systems engineering provides an effective methodology to create problem solutions for both known and unknown problems. Systems engineering is an organizational process, a management process and a psychological process with a wide variety of tools that yield effective results in a short time. Within the systems engineering process we organize and use tools and methods to: 1) turn visionary words into design requirements that reflect what the customer wants and needs, 2) create ideas for the concepts that will satisfy these requirements, 3) select the most viable concepts, and 4) choose the concept to take to the detailed development phase. If the process is haphazard, we may fail to satisfy the customer or, worse yet, have a very competitive concept but fail to get it to the customer in time. Effective system design is all about using time and effort wisely to create effective, innovative systems that satisfy or even astonish the customer. Systems engineering uses a wide range of tools such as: affinity methods, weighted objectives, Quality Function Deployment (QFD), Pugh’s method, morphological charts, functional decomposition and benchmarking. From a combination of these tools comes a clear definition and understanding of customer needs, together with the ability to turn a few inspiring “capabilities statement” words into a statement that resembles a homework problem. Unfortunately, the AAE curriculum does not include requirements that students become acquainted with these tools or the systems engineering process. It is left to individual instructors to insert this into their own courses. In the university, all courses involve projects, since even a standard homework assignment is a project. Students begin projects at various stages. For a homework assignment, the customer is the professor and the product is an answer: for example, the stress in a given component due to a given applied load. There is only one correct product, although many wrong answers may be produced. The design process involves finding the correct assemblage of formulae that, when strung together, will create the correct product. If only product design were that easy! The university design experience depends upon where the design team enters the process. For instance, a national design competition has a very well-developed Request for Proposal (RFP)


Final Report xvii

with metrics such as maximum aircraft range or time aloft specified, subject to a weight constraint. The customer is the competition organizer and most of the main product objectives are clearly laid out with customer evaluation criteria clearly and measurably expressed in numerical form. The substantial initial effort that led the competition sponsor to select these numbers and the contest goals is hidden from the competitors. Design classes pick and choose product design launch points between Phase 0 and Phase 1. The point at which to begin is the choice of the instructor and has a cultural history peculiar to the instructor’s background, experience and values. The most difficult design problems begin with a fuzzy opportunity statement like “design a system which can stay aloft for a long time and locate lost humans.” Turning these words into a systems document that has as its outcome a feasible design with value to the customer is a major task. It involves questioning the customer and finding out what he or she really wants. Unless challenged, the customer may not have thought about the answers. The design project goals and background This semester’s class project to challenge students involves a modular, reconfigurable space system, sometimes called a “fractionated” space system. This system is launched in segments and as a result is upgradeable and repairable. This effort is described in the student’s final report. The origins of this type of space system trace back to a proposed system called TechSat21 which was intended to be an Air Force Research Laboratory technology demonstrator. The TechSat21 concept envisioned a group of collaborating satellites coordinated to function as a sparse-aperture radar to examine Earth based targets. This concept required formation flying and communications among all satellites. During my time at DARPA we examined similar systems, originally called Distributed Space Systems, to see if they could be used for commercial and military uses other than radar. The result was a resounding “yes.”4 At issue is the use of a system that would require modularity rater than monolithic satellite operation. This effort is still embryonic and as such is an excellent topic for a design class. Reconfigurable space systems involve a collection of similar and/or dissimilar cooperating elements whose capability grows as more elements are added. These elements may have centralized or distributed “team” intelligence and capability, with different elements reconfiguring or morphing as operational situations require. The modular elements can attach and detach and operate cooperatively while connected or disconnected, collaborating with each other and with other systems to adapt to changing situations and allowing a system to be repaired or to age gracefully. Space system reconfiguration includes changing functional, performance, or physical characteristics at will and doing it again and again when required. Reconfigurable systems require components that change features on demand to give the system specific capabilities which can be transformed or modified by a number of reversible actions to provide different

4 T. A. Weisshaar and I. Bekey, “Responsive, Multi-function morphing Space Systems,” AIAA Paper 2005-2005 AIAA LA-Orange County 3rd Responsive Space Conference, April 2005.


Final Report xviii

•Single-mission dedicated spacecraft•Inaccessible and expendable•Non-serviceable•Non-upgradeable•Non-refuelable•Fixed architecture: deploy only•Incapable of cooperating with other s/c•Dependent on ground control•Require dedicated large launch vehicles

•Multiple specialized systems needed•Poor interoperability•Force-fit system-of-systems•Extremely expensive, individually•Obscenely expensive collectively•Specialized spacecraft by each Agency•Not adaptable to changed needs•Cannot respond to unpredictable events•Very long time to deploy•Need large ground control support•Poor survivability

Results in:

•Single-mission dedicated spacecraft•Inaccessible and expendable•Non-serviceable•Non-upgradeable•Non-refuelable•Fixed architecture: deploy only•Incapable of cooperating with other s/c•Dependent on ground control•Require dedicated large launch vehicles

•Multiple specialized systems needed•Poor interoperability•Force-fit system-of-systems•Extremely expensive, individually•Obscenely expensive collectively•Specialized spacecraft by each Agency•Not adaptable to changed needs•Cannot respond to unpredictable events•Very long time to deploy•Need large ground control support•Poor survivability

Results in:

Figure 3 - Current space operation features and their effects

•Capable of performing multiple missions

•Reconfigurable/morphing space systems

•Semi-autonomous and adaptive systems

•Intelligent self-assembly from modular elements

•Serviceable, refuelable, upgreadable on orbit

•Requiring only high level ground commands

•Can serve both infrastructure and operations

•Modules can use smaller launch vehicles

•Adaptive mission capabilities

•Responsive to operational need changes

•Responsive to unanticipated mission needs

•Much shorter time to deploy

•Much greater survivability

•Inherently able to reconstitute capability

•Inherent interoperability

•Much lower cost for any capability

•Very much lower cost for collective missions

•Infrastructure supports all Agency needs

Results in:

•Capable of performing multiple missions

•Reconfigurable/morphing space systems

•Semi-autonomous and adaptive systems

•Intelligent self-assembly from modular elements

•Serviceable, refuelable, upgreadable on orbit

•Requiring only high level ground commands

•Can serve both infrastructure and operations

•Modules can use smaller launch vehicles

•Adaptive mission capabilities

•Responsive to operational need changes

•Responsive to unanticipated mission needs

•Much shorter time to deploy

•Much greater survivability

•Inherently able to reconstitute capability

•Inherent interoperability

•Much lower cost for any capability

•Very much lower cost for collective missions

•Infrastructure supports all Agency needs

Results in:

Figure 2 - Multi-functional space system capabilities for transformational operations

capabilities quickly. This may require that elements of the system separate and re-join many times. Examples of future, military multi-functional space systems include:

• A space system with principal functions of Intelligence, Surveillance, and Reconnaissance (ISR) that can defend itself with non-lethal or lethal responses when attacked

• A space system with ISR observation functions that can trade performance for survivability by separating and dispersing its elements to react to a threat, then regroup them when the threat subsides

• A multi-spectral ISR system that can become an electromagnetic air or surface attack weapon when called upon, then resume its ISR function

• A communications system that morphs into an active jamming/EW weapon when needed, and then resumes its communications functions

As indicated in Figure 2, today’s space systems are very limited, very expensive and, once in orbit, are non-repairable and unresponsive to emerging technology changes. Space operations are expensive because of the large costs associated with launch, the constant and detailed “care and feeding” the systems require, and the fact that the systems—once in orbit—cannot be repaired; they must be replaced if any critical component failure occurs or if an upgraded capability is required. With a few notable non-DOD exceptions such as the Hubble Space Telescope, current space systems are


Final Report xix

Fuel and spares depotFuel and spares depot

Focal PlaneFocal Plane

Data ProcessingData Processing

Comm linkComm link

Power relay / Sensor suitePower relay / Sensor suite

SunshieldSunshield

Fuel and spares depotFuel and spares depot

Focal PlaneFocal Plane

Data ProcessingData Processing

Comm linkComm link

Power relay / Sensor suitePower relay / Sensor suite

SunshieldSunshield

Figure 4 – Fractionated space telescope concept

inaccessible and therefore non-serviceable, non-upgradeable, non-refuelable—and, as a result, expendable. This is NOT what space systems should be nor how they should operate. The future provides daunting challenges to both the military and commercial establishments. For instance, the collection of “non-State” actors, including terrorists, multi-national movements, trans-national industries, as well as numerous “peer” competitors, will make this a very different world in which to live and operate space systems. Non-U.S. entities, such as SSTL (Surrey Satellite Technology Limited) in England is a first class satellite educational and development center. It sells satellites and space technologies to everyone. An on-orbit capability and technology insertion through SSTL is happening at a rapid rate. There were fourteen non-U.S. micro-satellites (in the 11-92 kg. class) launched in 2002 (including Brazil, China and Algeria). Future systems must be multi-functional to cope with changes in the operational environment. Figure 3 summarizes the advantages of multi-functional space systems of the type developed in AAE450. Some needs are easily recognized a priori, and are included in the original design, but other critical needs may arise only after major systems are in orbit or have been used in situations for which operational learning is important. The Hubble Space Telescope is perhaps the best and worst example of the value of a modular system. Hubble was launched in 1986 and immediately ran into unanticipated problems that affected the imaging capability. A visit by astronauts repaired the system and essentially recovered a failed satellite system. Next year, Hubble will again be repaired by astronauts, at the expense of a $900 million dollar effort. The James Webb Space Telescope will replace Hubble with a high tech solution but because to the extremely high orbit in which it will operate, James Webb will not be repairable by astronauts. To take advantage of system modularity, Bekey5 suggests configurations like the one shown in Figure 4. This concept has few physical attachments and substitutes wireless communication and power links for the usual bus structure. While the technology needs are daunting, the advantages are apparent. This system does not need to be launched as a single unit. As the result, risk is spread over several launches. In addition, replacement of a critical unit can be

quickly and inexpensively completed if modularity is considered in the original system architecture. Many new technology developments are required to make these systems possible. The technologies below are only a few of those required. Maneuver, persistence, disperseability, upgradeability, manipulation

• Metrology and control systems to enable hard docking or close

5 Ivan Bekey, Advanced Space System Concepts and Technologies, The Aerospace Press, El Segundo, California, 2003.


Final Report xx

precision formation flying, as well as component dispersal for self-protection • Ultra-precise and responsive formation flying algorithms for both close proximity and far

dispersal, and re-joining again • Docking techniques and equipment with adaptive, intelligent, autonomous interfaces, for

mechanical, thermal, electrical, and information transfer docked and separated • Manipulating, grappling, and attachment techniques • Proximity operations and knowledge sufficient to operate autonomously in both friendly

and hostile environments • Means to store parts and debris in a “boneyard” element that is part of the space system

Global situational awareness, autonomous mission-responsive action. Self-assessment • Mission-adaptive ISR sensors capable of variable search, zoom, track, and other modes • Shape-changing, large RF and optical reflectors • Adaptable electronics and algorithms to assess situations and change the space system

autonomously • Modularity to allow servicing, upgrade and exchange on orbit via simple module

exchange, with backward compatibility • Sensors and computation for status assessment, health management, and damage

assessment/control

Propulsion and power, persistence, sustainability • Power generation, storage, and distribution across modules, docked or not, with

capabilities adaptable to differing needs on orbit • Thermal management of individual units and the whole space system which are adaptable

to different configurations and needs • Efficient, clean, and unobtrusive maneuver propulsion for both attitude control and

translations—large and small—for modules and master spacecraft • Means for large, frequent orbit changes, especially plane changes, with little propellant

expenditure Unfortunately, many of these technologies, even if developed to maturity by themselves, must be tested together with other new technologies operating as an integrated system. None of these technologies are anywhere near that point. No integrated analysis, let alone testing, has yet occurred or even been proposed due to the infancy of the reconfigurable space system concept. Bekey provides the following recommendation. “The concept of … reconfigurable/morphing (modular) space systems is so new that industry must … define new concepts and prepare demonstration proposals. This must include definition of the desired operational capabilities with the various new technologies working together. This is as difficult as the definition of the technologies themselves, if not more so, but crucial to obtaining the desired operational responsiveness.” (Italics mine.) The Mission Opportunity Statement – Version 1 The original mission opportunity statement distributed to the class early in the semester anchored the effort to a reconfigurable modular system with a single task, viewing the Earth. In part, it read as follows:


Final Report xxi

Terrestrial-imaging and space-imaging systems, such as the Hubble Space Telescope produce high value data. The resolution and image quality of these systems are limited by optics design. The optical requirements—lens and mirror sizes in particular—determine critical system design features such as system weight which, in turn, determines orbit selection and selection of launch vehicles. This has an important effect on system cost. Component miniaturization, as well as innovative design approaches, make it likely that space system weight and volume can be reduced during the next decade, resulting in improved image quality at reduced cost. Technological innovations include development of sparse-aperture mirror concepts, actively controlled membrane mirrors and modularized space system concepts launched on small rockets. System concepts using virtual links between modules enabled by wireless systems and adaptive control algorithms will reduce structural component weight. Space systems are very costly. The cost of delivery to orbit is a significant part of the total budget. A requirement to use a specific launch system with restricted weight and a defined payload fairing can impose limits on the overall system length, diameter and mass. The life of successful systems is short due to demands such as propellant for station-keeping and orbital changes. Critical system components are often designed to be redundant, adding to weight and cost. This Design Opportunity is looking for innovative concepts for a new, space-based Earth observation system (optical) to monitor weather over a large area, but also have the ability to focus on specific smaller areas of interest. This capability will provide both storm monitoring such as hurricanes, cyclonic activity and damage assessment. This space system will also have the ability to be converted into a military use during future international crises and turn its attention to space surveillance in nearby orbits. As a result, the system must have self-protection features.

I believe that students are important stakeholders in the design class. This mission opportunity was discussed extensively. Several concepts were developed. Despite my fondness for the project, I sensed that this mission was not ardently embraced by the majority of the class. They saw the mission as something that was already being done, even though I had added a new feature or two to the mission opportunity. As a result, we spent two more weeks discussing new missions and finally developed the mission that appears in their final report. The difference between the two missions is the requirement for multi-functionality. Both an Earth-viewing telescope and a space-viewing telescope were required. In the end, the final mission, developed with substantial input from the class, just as it would be in real life, was a better class mission. It was also more challenging.

Lessons learned Professors, no matter how old, are never done learning about how to create and manage experiences that improve the intellectual climate and also provide valuable training to students. Successes and disasters abound because undergraduate students come in a wide variety or


Final Report xxii

intellectual and academic shapes and sizes. The structure, methods and the demands of a design course do not fit all, challenge all or satisfy all. As a result, it is well to summarize lessons learned to create a record from which others, including other professors, can learn. First of all, let me say that it was a pleasure to supervise, lead, teach and interact with all, not most, but all students in this class. With one exception, these students were 5th year seniors and had a wide variety of experiences to share. All had a strong work ethic, although some were sleep-deprived and cured the deprivation during their colleagues’ presentations. Geoff and I conducted an exit survey and had a face-to-face with 13 of the 15 students in the class. Each class member was asked to rank his or her team members in order of value to the team and other team member characteristics such as ability to contribute, timeliness and respect for others. In a few cases these peer evaluations uncovered prime contributors whose value to the project had been overlooked by “the management.” The face-to-face exit interviews produced feedback the peer comments—together with ours—and identified areas where the process could be improved. All but two class members participated. These interviews were cordial, valuable and took two days to complete. When reading the student design report that follows, you should be aware of lessons learned, student attitudes and professorial successes and failures. The subject of technical writing came up more than once in the exit interviews. I stress the importance of technical writing and use the design class to have students practice skills to develop concise, readable reports. “How to” material is distributed in class. Assignments are graded and comments are given, sometimes in what is perceived as a brutal manner. These assignments form a portion of the student’s final grade. Several students objected to the five writing assignments distributed during the semester. One felt that the writing assignments “got in the way” of design activities. Another felt that the writing assignments were “make work” exercises that were “sprung on them at the last minute.” Of course I disagree, but one cannot sweep away perceptions with logic. What I should have done, and what all instructors must do, is to explain that there will be X writing assignments. The purposes of the assignments are threefold: 1) to improve writing skills, 2) to surreptitiously provide direction and guidance for the final report by having the students address topics pertinent to the final report, and 3) to clarify that the student understands design principles and is making progress on the design project. When the instructor does not believe that one or both of the latter objectives are being met, assignments are “sprung” and “progress” is interrupted. A second area of dissatisfaction was that students were prevented from “beginning the design project” because we took time to develop the mission statement, develop system requirements and learn about structured approaches before being allowed to begin the intensive, time-consuming job of synthesis and sub-system definition. This objection, voiced by at least three students, was expressed with what might be approximated as religious fervor.


Final Report xxiii

Again, the “fault” is mine. My approach assumes that the class understands what design and synthesis involves and how much detail occurs at the front end. The remedy is better preparation on my part. Even with this preparation, there were some strongly held beliefs about how design activities are organized and pursued. In a few cases, I doubt whether or not I would have been successful in changing minds and convincing students that I knew what was going on in the “real world.” The issue is what the term “design” means and the scope of design. Closely related to impatience with “beginning the design” as a source of dissatisfaction is the application of a “systems approach.” Everyone knows what “systems engineering” is—or so it seems—until a discussion of how to conduct a systems engineering approach to design. At its highest level, a systems approach will provide the high level “what’s” for the system, not the “how’s.” As the design progresses, this list will change slightly. It is curious that in a curriculum that is as multi-disciplinary as aerospace engineering, there is no formal, required introduction to systems engineering. There are at least five textbooks that cover this subject and how to apply it to design. Material from several of these sources is distributed early in the semester. A systems approach, beginning at the beginning, is essential to the success of a design effort so that students work hard, but also work smart. I need to do a better job developing a “mini-course” on this approach and relating its value to the project. Included in the systems engineering approach is the use of tools to guide the design efforts so that students work smart. There is no joy in working long hours on a project only to discover that your team has a solution to the wrong problem—and that your competitor has a better solution to the “right” problem. One of these tools is functional analysis in which the team is asked to break the system down into components that “provide” a purpose in response to a system need or operational objective. If the process is done efficiently and thoroughly, a team is led rapidly to the component design and integration effort in which trades are made and components selected. It is here that time is recovered from the earlier definition activities that students object to. Then, the component requirements begin to appear in terms of materials specifications, performance specifications and details that are traceable to higher level requirements that interpret the customer’s sometimes fuzzy objectives. It is fair to say that the team members who used the systems approach were confident that they knew what was required and, in some cases needed no help from me. Please read Chapter 3 to see the team’s view of how to apply this approach. Finally team dynamics were a source of extensive discussion, both pleasant and unpleasant. Reading the student evaluations of their fellow team members, I was impressed by the candor and fairness of the comments. There is always a chance that a student, when asked for an evaluation, will use it to settle a score or aim at retribution for some perceived slight or mistreatment. This did not happen even once. The most creative ideas are generated by individuals; the best designs are created by teams. Whatever, the management style of team leaders, it is imperative that everyone have or feel that


Final Report xxiv

they have a voice in design decisions. No one should be marginalized or feel marginalized. Since the focus of AAE450 activities is in conceptual design, the success of a team effort depends on team size. Smaller is better. Hindsight is usually better than foresight. If I were to go back to the beginning of the semester, I would divide the class into two teams, one with eight members and one with seven members. This would have had the effect of requiring more multi-disciplinary tasks for each team member, but would have also had the result that two competing designs would be brought forward. Teams would challenge each other. The key here is to develop an assessment tool so that the instructor does not ask for someone to do something outside his or her capabilities. However, this is Purdue and I refuse to believe that this is a problem. In the final analysis, the course could have been better, we could have done more, we could have done it differently, I could have been more effective, and I could have been more perceptive. However, as you read the student report remember that—to a man or woman—the team tried very hard, developed camaraderie, did what they thought was right, had a great deal of pride in what they did and ultimately learned from this experience. That is what education is all about. The students met the challenges we provided to them and even provided more challenges that we had not thought of. They deserve to be proud of themselves as individuals and as a team. Both Geoff and I wish them well in their future careers.


Final Report 1

1.0 – Background: Future Needs In today’s society we throw away small appliances, such as microwaves and toasters that are, “too expensive to fix”. This mentality has carried over to the aerospace industry as satellites are considered “too expensive to fix” and are simply turned off, or deorbited.[1.1] This has become so typical that most satellites are designed never to be fixed, upgraded or repaired. Satellite systems such as Hubble, the Global Positioning System (GPS), and numerous spy satellites are nearing the end of their lives awaiting replacement by new and better systems. 1.1 – New System Requirements A look into the future shows that traditional satellites will become obsolete and unable to meet the basic requirements of a changing world. The most desirable features of future space systems will be the responsiveness of the system and multi-mission capability. The ability to respond to a situation within hours and be able to perform several different jobs will be highly sought after. There are many potential future missions for spacecraft in both the civilian and military sectors. For civilian applications, maneuverable plane changes and being reconfigured on orbit may be necessary to obtain up to date weather photos or to replace a broken satellite. In the future, warfare is likely to move into space, providing opportunities for military applications. Weisshaar and Bekey suggest missions that will surpass today’s passive spy satellites; aggressive confrontation calls for evasion techniques such as instant fractionization to improve survivability. Such a scenario calls for weapons to identify and destroy enemy satellites, and the ability to repair damage incurred.[1.2]

It is difficult to meet the future needs of the spacecraft industry with monolithic spacecraft that operate as an indivisible unit. To meet these changing requirements a new approach to spacecraft design is necessary. Reevaluation of the design process can lead to the rediscovery and exploitation of previously impractical options. To pursue the idea of a multi-functional, fast-response satellite, technologies such as on-orbit docking, remote power transfer, component miniaturization, and on-orbit refueling will have to be developed. 1.2 – Fractionated Systems A fractionated space system is a new approach to meet requirements for future spacecraft missions. The concept of fractionated space systems involves breaking a spacecraft down into physically separated functional elements. Mathieu describes the concept of a “fractionated spacecraft” as similar to a monolithic system only with physically separated free-floating elements. The same subsystems of the typical monolithic bus would form an infrastructure of separated modules which operate together as a single system.[1.3] Fractionated systems are capable of reconfiguration because infrastructure components can associate freely with payloads as needed, and transit individually in smaller units. Figure 1 compares a traditional monolithic system to a fractionated one, illustrating the concept of how the subsystems of a monolithic system would work together in a fractionated system.


Final Report 2

Figure 1.1: Traditional vs. Fractionated Spacecraft [1.3]

Spacecraft components of a fractionated system need only to be able exchange energy, which can be classed as power and information, without resorting to conventional solid structures to connect them. Because the space environment applies no structural loads comparable to an Earth environment, no comparable structure need exist.[1.4]

Fractionated systems would be more appealing than monolithic options to customers for several reasons. They have the potential to solve the responsiveness situation and perform other functions such as the repair and upgrade of existing systems while replicating benefits from other concepts. They may include shorter delays between design concept and delivery.3 Additions to a system can be added without replacing the existing structures in a fractionated system. Because of this, fractionated systems may be a more cost effective route to space, excluding start-up fees for the initial infrastructure. The concept of fractionation could be applied to the deployment of a large monolithic satellite that exceeds available launch vehicle dimensions. A fractionated system could be launched in pieces and then assembled on-orbit to complete the monolithic system. Another possible application is one that is designed to have self defense and collision avoidance capabilities. A system that could separate and disperse itself to avoid an attack or collision with another object would be beneficial. With space becoming “crowded” and the potential for space based weaponry increasing daily, a self-defending satellite would be invaluable. This type of system could break apart into multiple pieces as an object approaches, then reassemble and continue operating with little service interruption, after the object passes. In the event of an attack on the satellite only pieces would be damaged, resulting in a system that is still functional.[1.2] 1.3 – Modular Systems A modular system is one that contains pre-defined modules that can be removed and replaced individually. In both modular and fractionated systems, payloads and infrastructures are placed into these separate replaceable packages. Being able to replace payloads allows the space system to vary its mission over time. In this way, fractionated systems serve as a subset of modular systems: both systems have the ability to interchange portions of their design. The physical connections present in a modular system are the primary difference between modular and fractionated systems. Despite this difference, both modular and fractionated systems are more


Final Report 3

adaptable than traditional monolithic systems. Interchangeable payloads increase the mission capabilities of modular and fractionated systems beyond those of a monolithic system.[1.1]

Modularity reduces the cost of spacecraft by introducing the idea of interchangeable parts that is being taken advantage of in projects such as the Joint Strike Fighter (JSF). With similar components throughout the system, development costs are lowered. Modularity also confines development and delivery costs to only the modules that must be changed or replaced when the mission of the system is changed.[1.5] Beyond affecting the cost, modularity expands the mission of a spacecraft: with the additional hardware and capabilities that modularity would bring to a spacecraft come more missions and increased usability. 1.4 – Mission Opportunity Statement

Purdue Advanced Research Projects Affiliates (PARPA) MISSION OPPORTUNITY STATEMENT

Modular, Fractionated Space System Advanced Concept Technology Demonstrator

1. Background Technology advances such as component miniaturization, as well as innovative approaches, make it likely that space system weight and volume can be reduced in the next decade, resulting in reduced cost. However, true cost reductions, coupled with new requirements for time reductions from concept to launch to operations make it unlikely that miniaturization will be able to reduce time and cost for some critical needs.6 New design concepts such as deployable actively controlled membranes, connected and unconnected sparse aperture mirrors, and modularized space system concepts launched on small rockets are among the many new concepts being proposed for future systems. In addition, virtual links between space system modules, enabled by wireless systems and adaptive control algorithms, will reduce structural component weight and blur the distinction between payload and the vehicle bus.7 Compared to terrestrial systems, space systems are very costly and have several disadvantages compared to terrestrial systems. First of all, the cost of delivery to orbit is a significant part of the system cost. The consequences of a single launch vehicle failure are enormous since the entire system will be lost, leading to long delays in the deployment of vital satellites.8 The life of successful systems is short due to demands for propellant to conduct station-keeping and navigation and control. Critical system components are often designed to be redundant, adding to weight and cost. This Design Opportunity is looking for innovative concepts for the prototype of a new, space-based Earth orbiting system to replace monolithic satellites. This new system, like a PC computer system, is to be versatile and modular so that it can be more reliable, more responsive, upgradeable and capable of being launched in component form so that risk of a single failure is 6 T. A. Weisshaar and I. Bekey, Responsive, Multi-function Morphing Space Systems, AIAA Paper 2005-2005, AIAA-LA Orange County Sections 3rd Responsive Space Conference, March 2005. 7 Ivan Bekey, Advanced Space System Concepts and Technologies: 2010-2013+, The Aerospace Press, El Segundo, California, 2003 8 Owen Brown


Final Report 4

removed. It should anticipate the need for large area terrestrial surveillance by having the capability of operating in different orbital planes without the need for large quantities of onboard propellant. Prime areas to consider are the area monitored and the resolution both in a wide angle and narrow field of view. It is important to understand how these choices influence orbit selection, system weight and complexity. 2. Opportunity Description Proposals are solicited for advanced concepts with the following features. The system’s primary mission is to replace the Hubble Telescope, but have additional capabilities. While Hubble looks only outward, this system must also look down at the Earth. How the system looks down, whether the additional system is attached to the main vehicle, the field of view and resolution are left to the bidder. The space system will have the capability of being launched in small vehicles, the vehicle type and number to be provided by the bidder, with special attention being paid to the system cost, responsiveness, assembly difficulty, ground check-out time and cost and the consequences of a failure of any one vehicle. The onboard processing units must be upgradeable and expandable at a reasonable cost. This capability will provide the system with the ability to change its mission by simply adding new modular features. The system must also be refuelable to extend the life of the satellite in low Earth orbit and to provide modest orbital parameter changes or fuel for assembly, as required. This system will have added complexity as a cost of versatility, however, the system must clearly show serious attempts to minimize complexity by simplifying or eliminating linkages and components. Ground communication will be centralized in West Lafayette, Indiana. The bidder will clearly show how the system is down-linked to West Lafayette and the number and location of intermediate telemetry. Communications and data rates will be determined by the bidder. Initial operation will occur in 2011, but a description of anticipated improvements in the system, based upon bidder knowledge of state-of-the-art development must be presented in the proposal. In addition, the bidder will present at least two additional Concepts of Operation, military or commercial, that can be supported by their design by the year 2021. 3. Data Requirements and Basis of Judging The final proposal must provide comprehensive engineering description of the design concept with detailed design information for major components and subsystems and their interfaces. This includes: 1. A detailed description of the system design and its operation. This includes professional

quality drawings of the system with key components and subsystems.

2. System requirements, requirement traceability and evaluation criteria used to develop the


Final Report 5

design.

3. A thorough description of the system architecture, together with the technologies and technical approach used to meet the mission requirements. This includes the degree of reconfigurability to allow the system to acquire new capabilities while on orbit or to acquire these capabilities while on the ground, during manufacture.

4. A review and description of all major concepts developed on the path to the final concept and a discussion of how features were excluded or retained.

5. Identification of critical technologies, relating them to problems and issues (resolved by trade studies or innovation) affecting system success. This includes a description of how technological advances in the future can be incorporated into the system in a cost effective manner.

6. Subsystem design details, including issues and assumptions, analyses performed, and tradeoffs considered for the system, including, but not limited to, the following:

a) launch vehicle features and modifications, orbital propulsion, attitude control and orbital maneuvering requirements, space propulsion.

b) mass properties and structural loads c) thermal system d) radiation shielding e) power and propellant f) communications and information processing

7. Test requirements for major systems

8. Estimated cost to develop the system, anticipated operational cost for the system and launches required.

4. Contacts All questions pertaining to this Mission Opportunity Statement should be directed to either Professor T.A. Weisshaar [email protected] or Geoff Wawrzyniak [email protected] (please copy one or the other) during class or after lecture. Key questions and answers will be posted on the class web site. This web site will also contain references and other information developed during the semester.

mailto:[email protected]


Final Report 6

1.5 – References [1.1] Feuchter, Christopher A., Charles A. Van Meter, Kurt M. Neuman, and Kalla J. Sparrow.

When is a Satellite Not a Toaster? 1991 Winter Simulation Conference, 1991, IEEE Computer Society. 3 Sept. 2006 <http://ieee.org>.

[1.2] Weisshaar, Terrence A., and Ivan Bekey. Operationally Responsive, Persistent Space

Systems. AIAA 3rd Responsive Space Conference, 25 Sept. 2005, AIAA. 4 Sept. 2006 <http://aiaa.org>.

[1.3] Mathieu, C, and A. L. Weigel. Assessing the Fractionated Spacecraft Concept. Space 2005,

1 Sept. 2005, AIAA. 15 Sept. 2006 <http://aiaa.org>. [1.4] Roberts, C, and P Eremenko. Cost-Benefit Analysis of a Notional Fractionated SATCOM

Architecture. Space 2006, 11 June 2006, AIAA. 10 Sept. 2006 <http://aiaa.org>. [1.5] Weck, O L., W D. Nadir, J G. Wong, G Bounova, and T M. Coffee. Modular Structures for

Manned Space Exploration: the Truncated Octahedron as a Building Block. 1st Space Exploration Conference: Continuing the Voyage of Discovery, 30 Jan. 2005, AIAA. 15 Oct. 2006 <http://aiaa.org>.


Final Report 7

2.0 – Design Approach and Requirements The Mission Opportunity Statement in section 1.4 of this paper presents a number of proposed requirements for the space system to be designed by the engineers of The Odyssey Project. To be determined is the overall concept of the spacecraft and how it will meet the mission requirements (Section 2.2). Identifying the design drivers of the system is essential; the spacecraft will be designed around these aspects of the system. The system requirements are backwards and forewords traceable to allow for each design aspect to be traced to its mission requirement. This system allows for the elimination of non-essential design aspects from the spacecraft design, reducing total mass and other burdens on the system. 2.1 – Space System Design The engineers of The Odyssey Project decided on a basic design of the spacecraft in order to perform the mission-essential tasks of space and Earth viewing (Please see section 3.0 for an in-depth description of the design process that brought the team to this final design). The basic concept is visualized in Figure 2.1:

Figure 2.1: Components of the Odyssey

To be noted in Figure 2.1 are the group-decided definitions of portions of the design created to ease the design process. The design is based around a “core” that carries power and data to the system. The octagonal core allows for the attachment of interchangeable “modules” on each of its eight sides, the modules aligned radially with the core. Module sizes vary, as will be


Final Report 8

explained in the section 7.5.3. The modules along with the core join to form a “segment.” Each segment will have a main function. There are three primary segments that are planned to be connected via interlocking cores: a propulsion segment; an optical segment; and an experimental segment. While the details of each of these segments will be expanded on later in this paper, they each perform their own missions: the propulsion segment keeps the spacecraft at altitude; the optical segment houses the space viewing and Earth viewing telescopes along with other essential system hardware such as solar panels to provide energy; and the experimental segment will house experiments from customers that will demonstrate future technologies. 2.2 – Determining requirements and Design Drivers While the Mission Opportunity Statement presents many possible requirements for a system, dividing these into tiers allows for prioritization when designing the system. Primary requirements of the system to perform the mission are placed in Tier 1, forming the basis for more detailed and design-specific requirements. A system driver is a system or component requirement which establishes the major features of the design. The satellite will look the way it looks due to these design drivers. These design drivers can be identified within the Tier requirements for The Odyssey Project. A small change in the features of a component that serves as a design driver will have major effects on the cost, schedule, or risk associated with the overall system.

2.2.1 – Primary System Requirements The Mission Opportunity Statement on which The Odyssey Project is based introduces several primary requirements for a variety of future spacecraft missions to be accomplished by The Odyssey. The first of these requirements is that the system must be an Earth orbiting system to replace monolithic satellites. This requirement allows for the incorporation modular or fractionated systems in the design. Actually, later in the statement it is mentioned that the system, “Is to be…modular.”[2.1] The modularity of the system will allow for it to be upgradeable. The need to accommodate new modules and upgrade of components serves as a design driver. The system would likely only have a simple unitary structure without these requirements. Instead, the Odyssey has numerous self-contained modules that dock to the side of the core, each drawing power from the core through the docking interfaces. Changing to a non-modular system would drastically change the design, resembling more of a monolithic system. Because the overall system is to be modular, all on-board systems must be broken down into components. As also mentioned in the Mission Opportunity Statement, the primary mission of the spacecraft is observation, both terrestrial and far field. The space viewing telescope is to be based on Hubble Space Telescope (HST). The formation of this primary mission is another design driver: without the space and Earth viewing missions, the spacecraft may have been designed for an entirely different mission such as communications. Such a mission may require a constellation rather than a unitary spacecraft, significantly changing the design of the system.


Final Report 9

Other teams such as orbits, guidance, structures, energy generation, and thermal must all accommodate the optical mission. For example, the large size and the large mass of the space viewing telescope required the design of a special “super module” to accommodate the system. This super module required redesigning the main core to avoid size conflicts. The decision to rotate the entire satellite to focus the space viewing telescope as opposed to a mechanism to point the mirror has affected the design of the orbit. Any drastic changes to the optics will cause other teams to modify their design leading to increased cost and a shift in scheduling checkpoints. There are several other primary system requirements mentioned in the Mission Opportunity Statement that may not serve as design drivers but are important to the system design. The system must be ready for launch by 2011, thus the call for a Technical Readiness Level of 7 for critical system hardware. The new spacecraft must demonstrate new technologies by 2021, hence the addition of an experimental segment on which these new technologies will be tested. The system must be controlled from West Lafayette, and as will be described in section 6.5.0, the Tracking and Data Relay Satellite System (TDRSS) was decided upon, relaying information to West Lafayette via the white sands ground complex (WSGC).

2.2.2 – Secondary System Requirements Secondary system requirements of the system include those that are implied by those mentioned in the previous section, yet not outright mentioned in the Mission Opportunity Statement. The ability of the system to provide power to modules is an example of such a requirement. The modules require the power in order to run system critical components, yet this is not explicitly stated as a primary requirement. The team overseeing the structural design of The Odyssey Project is greatly affected by secondary system requirements that are not always obvious. The need for the system to be modular requires a system to communicate to modules for docking purposes. This requirement calls for the development of a communications package that will be discussed in section 7.6.0. Since the system is supposed to demonstrate future technologies by 2021, the life expectancy of the system is expected to extend past 2021. This calls for structural analysis and special coating of the structures as discussed in section 7.0.0. Other examples of secondary system requirements include the requirement of the spacecraft to operate in Low Earth Orbit (LEO). While not initially stated in the Mission Opportunity Statement, this was later clarified by the customer, introducing the point that the customer is a prime source of non-obvious secondary system requirements. The requirement that the program must demonstrate overall cost savings over existing competitors is an implied requirement in most modern day design programs. However, with the enormous costs of space systems, this requirement is validly suggested. To ensure a fair cost analysis of the project, cost shall be calculated two ways and compared so as to minimize cost in all possible ways.

2.2.3 – Tertiary System Requirements Tertiary system requirements include those that are not directly necessary to completing the mission of The Odyssey Project, however it would benefit the program as a whole. The ability


Final Report 10

of the system to not only demonstrate future technologies but to implement those into system design in the future is an example of a tertiary requirement. Future hopes for The Odyssey include the ability to demonstrate a fractionated system, or components of one, through its technology demonstration capabilities. This is explained further in Section 13 which describes possible technology demonstrations aboard the experimental segment. 2.3 – Requirements This portion of the paper describes the actual System Requirements as defined by the engineers of The Odyssey Program design team. They are separated into a tier system, Tier 1 holding highest and most basic priority to accomplish the mission of the system. As the tier number increases, the detail of the requirement increases and is traced back to the tier above it.

2.3.1 – Explanation of Traceability The tier requirements as listed in section 3.2.1 can be interpreted in the following manner: Tier 1’s are the most basic requirements in order to fulfill the mission. An example would be labeled “1.” A more detailed requirement that stems from Requirement 1 that is not necessarily mentioned in the Mission Opportunity Statement but the design engineers is essential for the proper implementation of the system would be labeled “1.1.” This would continue with a “1.1.1” requirement that feeds from 1.1. This relationship is visually depicted in the requirement tree below: 1 Tier requirement 1.1 Tier requirement 1.1.1 Tier requirement 1.2 Tier requirement 1.2.1 Tier requirement 1.2.2 Tier requirement

2.3.2 – Tier Requirements Tier 1 1 The system propulsion system shall provide for altitude maintenance and attitude control. 2 The system shall be modular. 3 The primary mission of the system shall be observation. 4 The system shall demonstrate new technologies. 5 The system shall be controlled from West Lafayette, IN. 6 The mission shall be ready for launch by the year 2011. 7 The life expectancy of the system shall extend beyond the year 2021. 8 The spacecraft shall operate in Low Earth Orbit. 9 The program shall demonstrate overall cost savings over existing competitors. 10 The system shall maintain allowable temperatures and remove waste heat. 11 The system shall be reliable. 12 The power system shall provide adequate power generation.


Final Report 11

Tier 2 1.1 Attitude maintenance and attitude control thrusters shall use hydrazine as a fuel and helium

for pressurization. 1.2 Altitude maintenance thruster(s) shall provide a ΔV of 1m/s/yr. 1.3 The system shall carry enough fuel to deorbit at any time. 1.4 The spacecraft shall carry propellant to unload reaction wheels 92 times per refueling

period. 1.5 The spacecraft shall carry propellant to rotate faster than the reaction wheel. 1.6 The Attitude Control System (ACS) and altitude maintenance subsystems shall utilize

common fuel storage. 1.7 The spacecraft shall be refueled annually. 1.8 The spacecraft shall have spare helium capacity and capability to refuel future free-floating

modules. 1.9 All fuel system plumbing shall have a minimum of dual redundancy for leak isolation. 2.1 The system shall be upgradeable. 2.2 The system shall be able to adapt to new missions. 2.3 The system shall be serviceable or repairable. 2.4 The system shall have the capability to launch in component form. 2.5 The segment propulsion system shall deliver a newly launched segment from the launch

vehicle to the Odyssey and then deorbit. 2.6 The super-module propulsion system shall deliver a newly launched segment from the

launch vehicle to the Odyssey and then deorbit. 2.7 The super-module propulsion system shall deliver a newly launched segment from the

launch vehicle to the Odyssey, detach from the new module, and extract and deobrit an old module.

3.1 The spacecraft shall be capable of Earth and space observation using two telescopes. 3.2 The system’s orbit shall be appropriate for both Earth and space imaging. 4.1 Modules shall include docked and/or undocked experiments. 4.2 New technology shall be incorporated into the system by 2021. 5.1 The system shall provide continuous communications coverage. 5.2 The system shall employ ground systems to monitor and control satellite operations. 6.1 All primary mission critical hardware shall have a Technology Readiness Level (TRL) of 7

or higher. 6.2 The system shall meet the required deadline for primary launch of 2011. 6.3 The system shall employ existing launch systems. 7.1 The system shall meet the required deadline for secondary launch of 2021. 8.1 The system shall maintain an altitude of 785 km. 9.1 The system shall use existing facilities and “off-the-shelf-technology” to reduce costs. 10.1 The propulsion system shall limit energy transfer. 10.2 The propellant shall stay above 2 ° C. 10.3 Modules shall heat electronics and remove excess heat. 10.4 The optical system shall have tempergradients oless than 1/10th of a degree from side to

side. 10.5 Heat shall be removed from the core of the system. 11.1 Component and system testing shall be required.


Final Report 12

12.1 The power system shall provide adequate storage of power. 12.2 The power system shall offer adequate means to regulate and distribute power. 12.3 The power system shall provide memory to keep alive system. Tier 3 1.2.1 Altitude maintenance propulsion shall be provided by 1 kW hydrazine arcjet(s). 1.6.1 The system shall utilize three interconnected spherical 800 psi hydrazine tanks for fuel

storage. 1.6.2 The system shall utilize three interconnected spherical 6000 psi Helium tanks for fuel

pressurization. 1.7.1 The Odyssey shall be refuelable by Orbital Express. 2.2.1 The system shall dock via physical connections and a data and power interface. 2.2.2 The system shall be reconfigurable on orbit. 2.3.1 The system shall eject segments’ and modules’ non-permanent interfaces. 3.1.1 The system shall employ the Fine Guidance Sensors (FGS) to obtain Hubble optical

capability of 0.05 arcsec. 3.1.2 The Earth viewing telescope shall have a resolution of 1 meter. 3.2.1 The system shall maintain a desired orientation. 3.2.2 The system shall provide complete coverage of Earth. 4.1.1 The system shall provide space accommodation for additional modules. 4.1.2 The system shall have wireless connections to computers. 4.1.3 The system shall align modules and segments for on-orbit rendezvous with the satellite. 4.2.1 The system shall implement Technology Demonstration. 5.1.1 The system shall employ two trancievers to communication within access time. 5.1.2 The system computers shall be distributed. 5.1.3 The system shall communicate to TDRSS. 5.2.1 The system shall provide data processing and storing capabilities. 6.2.1 The launch site and vehicle availability shall be ensured. 6.2.2 The manufacturing capabilities of the system shall be ensured. 6.3.1 The launch vehicle shall insert the satellite into its desired orbit. 6.3.2 The launch vehicle shall provide a ΔV of 707 m/s. 6.3.3 The launch site shall meet orbital requirements. 6.3.4 The components to be launched shall fit in a fairing no larger than 4.6 m in diameter. 7.2.1 The future launch site and vehicle availability shall be ensured. 7.2.2 The future manufacturing capabilities of the system shall be ensured. 8.1.1 The system shall provide a ΔV budget of TBD. 9.1.1 The use of experimental materials and components to technology demonstrators shall be

limited. 9.1.2 Materials shall be selected for lightest weight and lowest cost. 10.1.1 Titanium standoff shall be used to separate the propulsion segment from other segments. 10.1.2 MLI shall be used to limit the heat absorbed from the environment. 10.2.1 Heaters shall have the capability to heat tanks, propellant lines, and catalyst point. 10.2.2 Low emittance material shall be used to limit heat loss. 10.3.1 Heaters shall be attached to electronics. 10.3.2 Radiators shall be placed on outer surfaces of the system. 10.4.1 The optical system shall be conductively isolated from the rest of the module.


Final Report 13

10.4.2 Small heaters shall be placed along the back of the optical mirror. 10.5.1 Heat shall be conductively sent to radiators at the exposed part of the core. 10.5.2 MLI shall limit heat to and from the modules. 11.1.1 The system shall survive launch loadings. 12.1.1 The power system shall provide fuse protection for equipment. 12.2.2 The power system shall provide emergency load shed. Tier 4 1.2.1.1 The system shall carry twice the required fuel per refueling period. 1.7.1.1 The system shall have an accessible refueling port for Orbital Express that extends aft

of the propulsion segment. 3.1.1.1 The system shall maintain a desired pointing accuracy of extensions. 3.1.1.2 The system shall employ a flexible mirror for space viewing. 3.1.2.1 The system shall only take Earth images during the daytime. 3.1.2.2 A two mirror reflecting system shall be employed for the Earth imaging system. 3.2.1.1 Gyroscopes shall be used to determine the vehicle’s attitude. 3.2.1.2 Reaction wheels shall be used to control the vehicle’s attitude. 3.2.2.1 The Earth telescope shall have a field of view wider than the greatest distance between

two ground tracks (113 km). 5.1.1.1 The system shall have a data transfer rate capability of 1.5 Mb/s. 5.1.2.1 The system’s computers shall have a processing speed of 2.2 GHz. 5.3.1.1 The system shall have an inclination accuracy of 0.1 degrees. 5.3.1.2 The system shall be able to have altitude accuracy of 10 km. 5.3.3.1 The launch site shall be capable of launching into a polar orbit of 98 degree inclination.


Final Report 14

2.4 – References [2.1] Weisshaar, Terrence A. "Modular, Fractionated Space System Advanced Concept

Technology Demonstrator." Purdue Advanced Research Projects Affiliates (PARPA) Mission Opportunity Statement. West Lafayette: Purdue University, 2006.


Final Report 15

3.0 – Design Concept Creating a design concept involved brainstorming concepts, selecting the top three concepts, and improving on them to decide on one. Pugh’s Method, a formal design method, was used to choose a final design concept. Pugh’s Method involves generating a list of criteria to compare each design. One design concept was deemed the datum. The other design concepts were compared to the datum. When comparing a design concept to the datum, the concept was either better (+1), worse (-1), or the same (0) for each criteria. Once the comparison was complete, the numbers were added to find the best design concept. [3.1] 3.1 – Brainstorming Design Concepts A design concept pool of fifteen designs was created by compiling each team member’s concept. Each design was assigned a number and placed in a packet with each design remaining anonymous as to the designer. Each team member used their own criteria to rank each design and placed votes for his/her top four designs. All of the votes were tallied leading to the top three design choices. These designs were separated from the rest of the designs by a significant margin (Table 3.1).

Table 3.1: AAE 450 Design Concept Voting Results 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 Hadi Ali 4 5 2 1 Nick Andrews 4 5 2 1 Craig Bittner 1 2 4 5 Matt Dennis 5 4 1 2 Jon Fromm 2 4 5 1 Elisabeth Hanssens 2 4 1 5 Matt Harvey 4 1 2 5 David Helderman 1 4 5 2 Norm Herbertz 4 5 2 1 April Miller 4 1 2 5 Chris Murphy 1 4 5 2 Ashley Ruic 2 4 5 1 Aaron Schinder 2 4 5 1 Nick Sochinski 4 5 2 1 Stephanie White 4 1 5 2 Totals 6 5 9 4 2 14 10 17 22 9 14 25 5 21 17 Rankings 11 12 9 14 15 6 8 4 2 9 6 1 12 3 4 # of First Place Votes 0 0 0 0 0 0 2 0 3 1 2 1 1 3 2 # of Second Place Votes 1 1 1 1 0 3 0 4 1 0 0 3 0 0 0 # of Third Place Votes 1 0 2 0 1 1 0 0 1 2 1 3 0 1 2 # of Fourth Place Votes 0 1 1 0 0 0 0 1 1 0 2 2 0 4 3 Total Number of Votes 2 2 4 1 1 4 2 5 6 3 5 9 1 8 7


Final Report 16

As the voting results showed the top three designs (#9, #12, and #14) have the most votes. Thus these three designs were chosen for further analysis and discussion and a more in depth look at the details of each satellite design. 3.2 – Concept Selection At the request of the user, the final three designs were renumbered for the next step of analysis. The conversion table (Table 3.2) is provided below to help eliminate confusion in later tables.

Table 3.2: Design Concept Renumbering Old Design # New Design #

9 1 12 2 14 3

The mission team broke up into sub-teams of power, propulsion/GNC, payload, thermal, and structural disciplines to further analyze each design concept. Each team presented their suggestions to modify concept design to meet the system and functionality requirements. Slight modifications were made to each design to help further describe it and to bring up any new design issues. Design #1 (Figure 3.1) is octagonal in shape and connects each segment axially in a traditional looking satellite system. The optical mirror is designed to be flexible to allow for different focal points and is placed on one end of the satellite. The propulsion system is placed on the opposite end allowing for thrust to be provided along the axis coincident with the center of mass, which is key for maneuvering. The system uses standard solar panels and should be able to be initially launched in two separate launches.

Figure 3.1: Design #1


Final Report 17

Design #2 (Figure 3.2) is spherical in shape and each hexagonal module attaches radially to the satellite. The primary optical system is attached as a series of modules on one end of the satellite, while the primary propulsion system is on the opposite end. The system uses a maneuvering arm to dock new modules with a circular docking ring. The system uses standard solar panels to provide power and contains all electronics internally to provide additional shielding.


Design #3 (Figure 3.3) is, in some ways, a combination of the first two designs. The main body is a hexagonal shape with additional segments connected axially similar to standard systems today. The main difference is that on this system the end segment is a half-sphere with numerous smaller docking ports for increased modularity. The primary optical system is attached on the side of the satellite and the propulsion system is on the end opposite the hemisphere docking module. This system also uses standard solar arrays and propulsion systems.



Final Report 18

Fifteen criteria were determined to evaluate each system using Pugh’s Method. The spherical satellite (Design #2) was selected as the datum point because it received the highest point total in the first round of voting.

Table 3.3: Pugh’s Method for Design Concept Criteria 1 2 3 New3

Innovative - DATUM S + Upgradeable S DATUM S +

Easily maneuvered (reorient) + DATUM + + Degree of modularity S DATUM S S

Multi-functional. Supports secondary missions S DATUM + + Easily stabilized - DATUM - -

Number & type of launches required + DATUM + + Ease of assembly + DATUM + + Optical adaptability S DATUM + + Ease of re-fueling S DATUM S S

Modular expandability + DATUM + + Orbital plane change capability S DATUM S S

Low drag S DATUM S S Critical components are high TRL + DATUM + +

Modules easily accessed - DATUM - + After the evaluation of Pugh’s Method (Table 3.3), Design #3 was the best choice of the three designs analyzed. Design #3 was then modified to incorporate design features of the other two designs to improve it. The final concept design was then evaluated again against the datum and determined to be sufficiently better than the former Design #3, the datum, and Design #1. Thus the final concept was New Design #3. 3.3 – Final Concept Description

Figure 3.4: Final Design Concept


Final Report 19

Figure 3.4 shows the final concept containing three segments: propulsion segment, optical segment, and bus. The propulsion segment contains the primary propulsion thrusters and electric propulsion orbital maintenance thrusters. It connects to the rest of the spacecraft through a central core of structure, docking ports, and wiring. In this way, the propulsion segment can be replaced. The optical segment has the same central core at its heart which is surrounded by mission critical modules (i.e. the telescopes). The bus segment is similar structure to the optical segment, but contains the hardware required to operate a space craft. Further experimental segments can be connected to this open end via the central core up to the point where the craft would become unstable due to its length. The core geometry is hexagonal. Later in the development process, due to excess space availability, the bus segment and optical segment were merged into one. One of the important features of this design is that the entire surface of the craft is used for housing modules. This idea originated from concept #2 except using a cylindrical rather than spherical form. This feature increases module capacity while maintaining symmetry and ease of control of the craft. Another important feature is the module housing structures. This feature aids in guiding the modules into place without limiting the size of the module. In this way, one module can be larger than another if necessary and still fit in the same slot without interfering with other modules. The final concept was a considerable improvement over the other three designs that were evaluated. The ability to remove each module individually without affecting other modules was one of the key factors of Design #2; this was incorporated into the final design by modifying segment connections. Overall this design had all of the features requested in the Mission Opportunity Statement and the ability to do more.


Final Report 20

3.4 – References [3.1] Weisshaar, Terrance. "Formal Design Methods." Purdue University. AAE 450 Lecture.

Purdue University, West Lafayette. 14 Sept. 2006.


Final Report 21

4.0 – Payload – Optics The optical capabilities of the Odyssey form the pinnacle of this space systems project. The Mission Opportunity Statement mentions the primary mission of the system as creating a replacement for the aging Hubble Space Telescope (HST).[4.1] Since its launch in 1990, Hubble has represented one of NASA’s few projects that gained a loyal public following. The high quality far-field images that document never-before-seen events such as the formation of stars (as seen in Figure 4.1 below) account for this popularity.

Figure 4.1: Hubble Image of Eagle Nebula

As Hubble ages, the need for a replacement becomes apparent. Incorporation of new optical technology to reduce mass and cost of a current space telescope comparable to HST created an interesting design problem for the optical team of The Odyssey Project.


Final Report 22

Equivalently important, Odyssey is expected to perform Earth viewing missions while orbiting in a 25 day sun-synchronous orbit. This orbit allots total terrestrial coverage in that period. To present a competitive Earth imaging product to the market, the system must be capable of a resolution comparable to others currently available. 4.1 – Scope and Purpose The resulting optics subsystem of The Odyssey Project was to be designed with the functional requirements mentioned in the Mission Opportunity Statement in mind. The design of this subsystem encompasses the imaging apparatuses chosen for both optical missions as well as of the spacecraft segment that houses them. Design decisions for the optical segment drove the design of the entire Odyssey Project due to the vital nature of its mission. Particular design drivers will be discussed in section 4.3. The desired result of this portion of the design is a better understanding of how both the Earth Viewing Telescope (EVT) and Space Viewing Telescopes (SVT) collect images, store and transfer the data and in turn complete the missions as stated. 4.2 – Tasks, Functions, Requirements and Design Methodology Referring back to section 2.3, the design requirements for the optics portion of the Odyssey is laid out and traceable back to the primary mission: observation of both Earth and space. As implied earlier, to present a competitive product, the system is to be comparable to HST for space viewing and be a competitive Earth viewing telescope, for example IKONOS (the satellite that obtains images for Yahoo Maps) and Quikbird. These requirements set the image resolution for the two subsystems to be 0.05 arcsec for the SVT and 1 m per pixel for the EVT. The pointing accuracy for both subsystems needs to be maintained to obtain these resolutions, calling for attitude control and precise pointing systems. The design of each of these subsystems as well as of the physical mirrored structure of each telescope is covered in forthcoming sections. Design methodology for the telescopes follows equations and techniques set out specifically for each subsystem. For example, the SVT requires long viewing times (on the order of hours to days) of the same point in space to get very clear and precise images. The EVT needs to have very short exposure times (hundredths of a second) to take clear pictures with minimal blur. The design overviews are in later sections of this report and necessary equations are found in Appendix I. 4.3 – Design Choices and Drivers To begin the design process of the optical segment, the engineers of The Odyssey Project had to develop a way, employing the idea of fractionated and modular systems, to accomplish two goals at once, not traditionally done simultaneously. After group brainstorming activities, two distinct categories emerged: two separated systems, each performing one of the two observations missions; or one physically connected system housing the hardware to perform both observation missions. After contemplation of aspects of the mission such as dynamics and maneuverability that would be affected by this design decision, a physically connected system was chosen. The choice of housing the space and Earth viewing hardware together were based on two main


Final Report 23

points: While most far-field space imaging is done far beyond the orbit of the Earth, the system that was to be compared to (HST) is in Low Earth Orbit (LEO), which is required by this system. At the same time, most Earth observation satellites are in LEO. If both subsystems are to be in LEO, physically connecting them eliminates the dynamic problems that would be encountered when maneuvering. Since the connection of these two subsystems was possible and it would potentially eliminate some orbital and maneuvering problems, it proved to be the more ideal of the two choices. Once a connected system was chosen, there was a choice between two mirror arrangements: a one-mirror assembly for both imaging missions; or two separate subsystems, one for Earth imaging and one for space imaging. A one-mirror assembly could perform two functions and would therefore lower the cost and overall weight of the system. A rotating faring would allow only one portion of the space craft to move, moving the imaging mirror assembly to focus on either Earth or a space object. More gyroscopes would be necessary to counteract the torques produced by the rotation of this one section to keep the satellite in equilibrium. The sectional rotation of the mirror while keeping electrical signal contact between segments would be possible with modern slip ring technology. Slip rings are metal rings that provide a continuous electrical connection through brushes on stationary contacts.[4.2]

The mass and cost penalties of the extra gyroscopes along with the increased power requirements for a one-mirrored system may counteract the advantage of a one-mirror system. The one-mirror subassembly would have to drastically change its focal length and image gathering techniques between Earth and space viewing missions. The cost of development of a dual functioning mirror assembly would rapidly approach that of two separate subassemblies. The reaction time between switching from Earth to space observation was also a concern. A one-mirrored assembly would limit the system capabilities, noting that both Earth and space viewing could not be performed simultaneously. Therefore, it was decided that two separate imaging subassemblies would be designed: one for Earth imaging and the other for space imaging. Since both of these optical subsystems were based upon already existing telescopes, their designs were based around existing hardware. The Fine Guidance Sensors (FGS) on HST, used to locate stars and position the satellite to obtain the required resolution, proved to be the most massive and largest component included in the design of the SVT. The FGS was chosen because it is the only system known that is capable of providing the desired resolution in LEO. The accuracy of the FGS is 0.003 arcsec, allowing for the 0.05 arcsec resolution of the overall mirror (the difference between these accounts for system errors such as mirror imperfections, etc). This drove the module design for the SVT subsystem, forcing the expansion of module length from the original 2 meters to 5 meters, and increased the number of one-meter-long modules that it takes up from 15 to 25. The choice to create this super-module for the SVT created the need for a second launch. The first launch would deliver the propulsion optics segments, minus the super-module. A second launch would deliver the SVT in the super-module to the empty space on the core of the optics segment. With all the changes made to the system to accommodate the FGS necessary for the SVT, the FGS can be placed as a system design driver.


Final Report 24

4.4 – Design Evolution The design for the Odyssey Project went through four levels of iterations before a final design was selected. Initial design configurations were based on proven concepts that have seen wide spread use in current functional space systems. Figure 4.2 demonstrates this will the first design concept illustration. The system was to be built in modular units on the ground, assembled into one monolithic unit and then launched into orbit. The design was to have three separate sections that would have either been launched together or placed into orbit individually. Each section would have had a set task, including the bus, propulsion and the experiment and optics sections. The optics section was deemed to be too exposed to perform accurate space observation. Also the design did not allow for any easy Earth viewing; the entire spacecraft would have to move to view the Earth.

Figure 4.2: Initial Design Configuration

The ideas from the idea shown in Figure 4.2 were edited and added to, along the concept to evolve into the next stage of design. Figure 4.3 shows the evolved design that was thought to have expanded the modularity of the satellite. This would have allowed very specialized modules to be included. One problem with this design was the spherical shape which would have limited volume for the optical mirror internally, so it would have to be mounted externally. This design also suffered from moments created by moving external equipment. Another problem was that the external mirror design would expose too much area to space. This design was rejected in favor of the third evolution which brought back a simpler approach to the design, as shown in Figure 4.4.


Final Report 25

Figure 4.3: Evolved Design

Design 3 decreased the complexity of the overall design and introduced the idea of a flexible mirror. It was initially theorized that one mirror could do the both Earth and space observations by changing the mirror’s focal length and radius of curvature. By incorporating this idea with a modular system, the design met the other requirements for the system. However, after calculations were done and working professionals in the optical field were consulted, it was decided that it would be impractical to use one mirror to do both observations.

Figure 4.4: Decreased Complexity Design


Final Report 26

These design problems led the team to the final design as seen in Figure 4.5: a hybrid of previous designs, except that there are separate optical systems for both Earth and space observation. The two telescopes are designed separately, each specific to its mission. Due to this feature, the designs of the mirrors do not sacrifice performance to complete both missions. The mirror assemblies are mounted in such a way so that both telescopes could theoretically be utilized simultaneously; assuming power and data rate transfer requirements are met. This design allows for the complete coverage of the Earth and the ability to image any desired point in space.

Figure 4.5: Final Design

4.4.1 – Earth Viewing Telescope To fulfill the requirements set forth in the Mission Opportunity Statement, the satellite must have the capability to view the entirety of Earth. The design of the EVT is the leading design factor necessary for obtaining the goal. This section details the process that created the final EVT for The Odyssey Project. The final design is comparable to Earth observation satellites currently in use today that produce high resolution images. The current standard for these crafts is a resolution of one meter. This means that one pixel on a digital picture is equivalent to 1 meter. These images can be found on popular websites such as Google® or Yahoo®. They are also able to do this cheaply and take pictures of any location on the planet. The design for the Odyssey’s EVT would allow for similar application.

4.4.1.1 – Design Decisions The final design of the EVT was obtained in two steps. Each step included design decisions and system requirements. It was also based on constraints set forth by other subsystems. In the following sections each major design decision for the EVT will be discussed.


Final Report 27

4.4.1.1.1 – Main Mirror The main driving factor in the design of the EVT was the ground resolution of the telescope. The engineers determined that a resolution that shows 1 m equivalent 1 pixel would allow the Odyssey to be a competitive Earth viewing system. This would make it competitive with current satellites in orbit such as IKONOS and Quikbird.[4.1] Computing the general size of the collector that is required for the resolution given is based on Equation 4.1.[4.3]

resolutionhD λ44.2

= (Equation 4.1)

D is the diameter of the main mirror, h is the height of the spacecraft and λ is the wavelength of the light. This was an important step in the design of the EVT because it showed how much space the mirror required. The mirror of the primary ended up being 0.976 meters in diameter.

4.4.1.1.2 – EVT Location Sensors need to be placed in proper locations to function correctly. In this respect, the EVT needs to be pointed towards the Earth to function as required. This means there are several possibilities so it must either point at the Earth, the telescope must move to point, or the craft orientates to point at the Earth. The location chosen for the EVT is the underside (Earth-facing side) of the Odyssey. This allows small movement of the telescope to point it at the target. This allows for the telescope to have short start-up and targeting times when an image site is coming into view.

4.4.1.1.3 – Mirror Technology Mirror technology has become advanced in recent years. New ideas like flexible mirror arrays and free-floating, self-inflating mirror arrays are being tested. The design required careful consideration. Flexible mirror arrays allow for the slight changing of the radius of curvature. of the mirror to change the amount of area it is fixated upon. A fixed mirror views the same area with the same clarity after launch. A flexible mirror can change how it sees an image location. This was not required by the RFP and the requirements do not require this change in focal length or area.[4.4] The free floating design was also not considered based on the requirement that all components must be attached to the spacecraft.[4.4] This left the third option of a static mirror. A static mirror is high TRL technology, one that has seen numerous uses on commercial and scientific observation platforms. It is an acceptable technology because the engineering work has been done by many companies. It has a proven track record of working in all orbits and on large scale telescopes, such as Hubble, and smaller satellites such as IKONOS. A static mirror costs a lot of money to produce. A single mirror costs roughly $20 million per square meter to create.[4.3] Also it is unable to change its focal length or curvature list more advanced mirror technologies. Also it is very heavy, glass has a density of 2200kg/m3, which makes it extremely heavy to launch a mirror of the necessary size into orbit.[4.5] See section 4.4.1.1.4 for more information on mirror material selection.


Final Report 28

4.4.1.1.4 – Mirror Material The primary and secondary mirror will be constructed of Ultra Low Expansion (ULE) glass. As Table 4.1 shows, the coefficient of thermal expansion (CTE) is the lowest for ULE glass. This is important when creating a mirror for space operation. If the temperature between the two sides of the mirror becomes too great the mirror will start to deform. This will cause the image to distort and produce pictures of lower quality. Using ULE this separation of temperature can be increased to allow for more leeway in designing a thermal system to support the mirror. Since glass is very brittle it needs to be reinforced to provide enough support to accomplish its task. This can be done either by designing a very thick but heavy mirror or by creating a supportive base. As Figure 4.6 shows, the mirror will have an internal aluminum truss to support the mirror. This reduces the amount of glass needed to have a fully supported mirror that can handle the stresses produced by launch and orbital insertion.[4.5]

Table 4.1: Properties of Current Mirror Materials[4.3] GPa g/cm3 ppm/K w/mK j/kg-K 1e-6 m2/s

ULE 68 2.20 <0.03 1.31 776 0.80Zerodur 91 2.53 <0.05 1.46 821 0.80silica 73 2.20 0.52 1.38 703 0.84SiC 466 3.05 2.37 300.00 660 146.00Borofloat 63 2.22 3.20 1.10 830 0.60

Figure 4.6: Model Mirror Design Construction[4.3]

4.4.1.1.5 – Visual or Infrared Another important design decision is the wavelength of light to be observed. According to Equation 4.1, the shorter the wavelength of light to be observed, the smaller the mirror is. If the inferred spectrum is to observed by the mirror it must be cooled to see the variations of heat released from the Earth. This requires cryonics to cool the mirror. This would increase the weight of the system by at least 300%.[4.5] The Sun-synchronies orbit allows the telescope to


Final Report 29

observe the Earth 70% of the time in daylight; an inferred system is not necessary, see section 5.4.1. Going to Near IR (0.8 µm) through IR (1.5 µm) triples the size of the mirror if just using visual light (0.5µm); this is based on Equation 4.1. It was determined due to space concerns to just stay in the looking the visual region.

4.4.1.1.6 – Image Collection The image collection process works by taking several small snapshots of the planet and then piecing those images together on the ground. This is done fort several reasons. Fist the image take by the EVT is not wide enough to collect the entire space between two ground tracks, the maximum distance between two ground tracks is 113 kilometers and the telescope can only take images of 10 kilometers in diameter (see section 4.4.1.3.1 for more information). This constraint of technology requires the system to point at several different areas on the ground to get the whole picture. Figure 4.7 shows how the telescope will take each picture to gain the full area between ground tracks. The images once taken are then stored on the spacecraft’s memory until the high speed communication bands are free to transmit each picture down to the ground station. Once on ground the images are then pieced together to form the entire image requested by the customer. The other way to take a series of ground images is to take the full area of the target between the ground tracks, this has several disadvantages. First the mirror must be larger to take more images per shot. Second the CCD chip must contain more chips to record the image, this increase in chips causes the memory storage requirement to increase per image, see section 6.6.2. Also when transmitting individual pictures the amount of bandwidth needed to transmit these images becomes exceeding larger then commercially available, see section 6.5.3. The first process was selected because of its lower requirements on storage and communication designs.


Final Report 30

Figure 4.7: Process of Image Collection

4.4.1.2 – Preliminary Design The preliminary design of the EVT was to have a telescope that combined the functions of both the stellar observation requirement and Earth viewing requirement. This arrangement was quickly scrapped because it is a design problem that is beyond the scope of what could be current technology is capable of producing. This caused the shift to two independent telescopes. The following sections of this report details the steps taken to produce a system that could meet the requirements set forth in the RFP.

4.4.1.2.1 – Trade Studies One of the trades is the number of mirrors required for the EVT. One mirror requires very long focal lengths and is very simple to contract. Multiple mirrors form smaller focal lengths, but require more complex truss arrangements to keep them from moving. The design was to minimize space, so multiple mirrors were the best choice. The final design has two mirrors to reduce size and weight with minimal complexity, and thus cost, to the final design.


Final Report 31

The mirror used by the EVT can be sized differently depending on the type of telescope being used as the primary collection system, see EVT appendix on the different types of telescopes used for Earth observation. A series of lens adds more mass then a traditional reflecting mirror arrangement. Because it would require lens of similar dimensions of the mirror placed in series to convert the image to size acceptable to fit on a properly sized CCD chipset. Each lens adds the weight of another primary mirror to the system. If the telescope should require three lenses to properly shape the image would be the equivalent of producing three mirrors. Additionally lens used for Earth observation must be perfectly clear because any distortion of the material would cause the image to be fuzzy or have spots.[4.3] Which these considerations having a two mirror assembly was optimal for this design.[4.4]

The image generated by the EVT can be created in two ways: create the image in one pass or take several different images and combine them later. The first way requires the telescope to have a large field of view that can take the entire path as the satellite travels over the ground track. This is an expensive way. To accomplish this, the telescope requires a large collection area and has a large instantaneous field of view. This reduces the amount of computational work that needs to be done, and could be transmitted directly to the ground. If the image is take over several sections it allows the telescope to be any size; except the image must be combined to create a target image.[4.4]

4.4.1.3 – Final Design The final design of the EVT encompasses all of the decisions made to create the optimal system for use with The Odyssey Project. The final design of the EVT takes 11000 x 11000 pixel images of a target area of 10 kilometers diameter. To be able to do this the telescope pivots to allow full coverage of the area needed. The telescope fits in standard size module which is attached at sections 2.2 and 2.3 of the optical segment. The telescope is a standard two mirror design, which collects an image on the primary mirror, reflects that image onto a secondary mirror and then onto a CCD chip which records the image. The entire mirror assembly is placed within a light reducing shroud that keeps the telescope operating at nominal temperatures and improves image quality. The following section details the final components that make up the final design of the EVT.


Final Report 32

Figure 4.8: EVT Cutaway

4.4.1.3.1 – Components and Specifications Figure 4.8 shows the design of the EVT. The EVT is composed of a mirror assembly, a light reducing shroud, a pivoting control system and module with connections. The mirror assembly holds two mirrors and a CCD chipset. Since the EVT cannot take the full view of the area it needs in one pass, it must move its pointing target to retrieve the full area (see Figure 4.7). The EVT can pivot within 0.05° of the target location this means a 61 meter off target image, but with the system overlapping each image it takes missing by that amount on a 10 km image is not a major design problem. Since the


Final Report 33

image collected by the EVT is 10 km in diameter the distance missed does not create a design problem.

The shroud is an object that reduces the amount of indirect light coming into the telescope. The shroud is composed of an aluminum structure for support and has a galvanized aluminum skin. The outer surface is painted white to reflect inferred and visual light away from the telescope. This helps keep the mirror at a lower temperature. The inner skin of the shroud is painted black to absorb light and has baffling, small grooves cut into the shroud to reduce reflected light. The sun-shade places on the end of the shroud places a direct barrier between the telescope and major light source, also reducing the amount of light coming into the telescope.

Figure 4.9: EVT Shroud

The mirror assembly lies inside the light reducing shroud. The mirror assembly’s main function is to provide a solid structural component to the mirrors. This will reduce the amount of drift viewed by the mirrors. Also it provides fixed lengths for each the components in the assembly, primary mirror, and secondary mirror and CCD chipset. These components need to be placed at fixed lengths or the images received by the telescope become out of focus.[4.1]


Final Report 34

Figure 4.10: Mirror Assembly

The CCD (Charge Coupled Device) chipset has been designed to capture a 10 kilometer diameter ground image. This requires from the system to have a diameter of 10,000 pixels on the chipset. Current CCD technology limits the minimum size of the CCDs. The current industry standard is 15μm per pixel. This means that we need have at least 150mm diameter for the chips to have the proper capture area. Figure 4.11 shows how the CCD layout, the chips to gain the maximum coverage area. The chipset incorporates 21 CCD chips each 2000 x 2000 pixels; each chip is 30 mm by 30 mm. This creates an area of 18900 pixels. This radial arrangement captures all of the light from the secondary mirror. To have fast shutter time for the image collection, the CCDs will be digitally shuttered. This improves on slower mechanical shutters, by telling the CCDs when to “look” the system responds faster allowing for faster response on ground tracks.


Final Report 35

Figure 4.11: CCD Chipset Layout

4.4.2 – Space Viewing The primary mission of The Odyssey Project is to design a replacement for Hubble. HST’s marked success in space exploration suggests it as a good starting point for the Odyssey design.

4.4.2.1 – Concept Design: Methodology and Trade Studies Referring to the Tier requirements found in section 2.3, the most basic and essential requirements that the design of the SVT addresses include the following:

3.0.0 The primary mission shall be observation. 3.1.0 The spacecraft shall be capable of Earth and space observation using two telescopes. 3.1.1 The system shall employ the Fine Guidance Sensors (FGS) to obtain Hubble optical capability of 0.05 arcsec. 3.1.1.1 The system shall maintain a desired pointing accuracy of extensions. 3.1.1.2 The system shall employ a flexible mirror for space viewing.

3.1.0 was resolved earlier in section 4.3, when it was decided that two optical subassemblies would be necessary to accomplish the mission. The two subassemblies are connected and are unable, due the preliminary decision, to separate in a fractionated way to operate independently.


Final Report 36

With these beginnings, as well as the metrics of HST around which the Odyssey was designed, a detailed design for the SVT can be determined. They key to this design is taking the basic conceptual ideas of the HST and incorporating new technology, forming an updated, more reliable and survivable version of Hubble. To begin the design process, a basic understanding of Hubble’s imaging capabilities was necessary. HST consists of a basic dual mirror assembly in which light enters the telescope and strikes a concave primary mirror, which acts like a lens to focus the light. Light from the primary mirror is reflected to a smaller secondary mirror in front of the primary mirror, then back through a hole in the primary to the main “camera” behind the focal plane. This dual mirror configuration is known as Ritchey-Chretien Cassegrain design. Hubble's angular resolution is 0.05 arcsecond. To understand the meaning of an arcsec, relate it to a full moon which extends approximately 0.5 degrees (1800 arcsec) across one’s view of the night sky.[4.9] This is the "sharpness" of Hubble's vision. A person that could see as well as Hubble could stand in New York City and distinguish two fireflies 1 meter apart in San Francisco.[4.10] The HST’s main camera, the Wide Field and Planetary Camera 2 (WFPC2) is made up of four postage-stamp-sized CCD (Charge Couple Device) chips. CCD’s are electronic circuits made of light-sensitive elements (pixels). Each of the four 800 by 800 pixel CCD’s that make up WFPC2 contains 640,000 pixels. The light collected by each pixel is translated into a number. These numbers are sent to ground-based computers, which convert them into the famous Hubble images.[4.11]

The first decision to be made on the Odyssey’s SVT was the mirror assembly layout. Two possibilities were considered: HST’s dual mirrored Ritchey-Chretien Cassegrain design, or a single mirror system. These two designs are easily distinguishable from each other: for a single mirrored system, the focal length, or the length at which the CCD assembly must be placed to record a focused image, extends outward from the mirror forming a tall truss as shown in Figure 4.12 A. The focal length and radius of curvature are able to be calculated by hand given the desired resolution. A design portraying the Ritchey-Chretien Cassegrain design, as shown in Figure 4.12 B employs a secondary mirror that places the focal point behind the primary mirror. Therefore, the CCD assembly is also placed behind the primary mirror. Ray tracing software is required to calculate the focal length of a double mirrored system. The height of aperture of a double mirrored system is less than that of a single mirrored system.[4.12]


Final Report 37

Figure 4.12: One Mirror (left) vs. Two Mirrors (right)

This being said, the two approaches were compared to see which would serve as a more viable option for the Odyssey’s SVM design. Since the focal point of a double mirrored system lies beneath the primary mirror, the mirror assembly must be raised a significant distance from the core to accommodate the CCD assembly and other imaging hardware. This would offset the center of mass and may cause it to be positioned outside of the core, causing instability. The lack of ray tracing software would lead to estimations of the distance between primary and secondary mirrors for this system. The errors associated with these estimations along with the instability problem caused by an offset from the core deterred the design team from choosing a dual-mirrored design. By default, the single mirror system was chosen as the SVM design. The next major design decision to be made was the material selection for the primary mirror. The mass of HST’s primary mirror is about 800 kg, proving that the material choice here could affect the total system mass considerably. Hubble employs an ultra-low expansion silica glass primary mirror coated with a thin layer of pure Aluminum to reflect visible light. A thinner layer of Magnesium Fluoride is applied to the Aluminum to prevent oxidation and to reflect ultraviolet light.[4.13] Because they are made with glass, Hubble's mirrors have to be kept at a nearly constant room temperature (about 70° Fahrenheit) to avoid warping.[4.14]

A segmented mirror made of Beryllium with a density of 25 kg/m2 and a 2.4 m diameter was calculated to have a mass of 113 kg. Another argument, other than mass conservation, for the use of beryllium instead of glass, is the price per square meter of each of these materials. A study by NASA and the Advanced Planning and Integration Office shows the cost per square meter of the glass mirror on HST is about $10 million while compared to less than $3 million per square meter for James Web Space Telescope (JWST, made of beryllium).[4.15] Both the HST and Odyssey primary mirrors have the same diameter of 2.4 m. Therefore a segmented Beryllium mirror is a less massive and more cost effective option as opposed to a solid glass mirror. Figure 4.13 shows the comparison between these two types of mirrors graphically.

CCD

CCD

Primary mirror

photons

Secondary mirror

Primary mirror

photons


Final Report 38

Cost and Mass ComparisonsHubble Solid Mirror vs. Segmented Beryllium Mirror

0

100

200

300

400

500

600

700

800

900

Hubble (solid) Segmented Beryllium

Mas

s (k

g) o

r Cos

t ($M

)MassCost

Figure 4.13: Cost and Mass Comparison of Mirrors

4.4.2.2 – Final Concept Selection The final concept design can be determined from the concept decisions described in the previous section. This section of the paper is split into the major components that make up the design of the Space Viewing Telescope.

4.4.2.2.1 – Segmented Beryllium Mirror Professor Finley, an optics professor in the Department of Physics at Purdue University, offered much insight in to the design of spacecraft optics and assisted calculating the vital metrics for the SVM, as depicted in Table 4.2.

Table 4.2: Descriptive Metrics of Odyssey’s SVT Mirror diameter (D) 2.4 m

12 cm/deg Plate scale 0.59

arcsec/pixel Mirror radius of curvature

13.74 m

Focal length (f ) 6.875m Resolution 0.05 arcsec Mass of mirror 113 kg Mass CCD structure 12 kg Mass of Al supports and reshaping mechanisms

12.7 kg

Total mass of SVM 137.7 kg Power req’d by SVM 1565 W


Final Report 39

Each of these metrics was based on the original adoption of 0.05 arcsec as the desired resolution of the SVM. This data along with the size of the CCD chip assembly (2cm square) allowed for the plate scale to be calculated. Plate scale is an optical industrial standard that relates the angular field of view of a frame to the size of the detector, either in terms of a dimension or per pixel.[4.16] Both versions are listed in Table 4.2 to give an accurate description of the Odyssey’s optical system. Note that the total SVT mass includes the mirror mass, CCD structural mass, the Aluminum supports as well as the reshaping mechanisms for the mirror masses. For descriptions of the required calculations to obtain these metrics, please refer to Appendix A.4.4.2.[4.12] The Beryllium mirror is made of hexagonal segments that will make up the 2.4 m diameter mirror. The segments will be positioned using micromotors and a wavefront sensor to obtain the correct radius of curvature once the SVT docks with the optical segment. This will act like an initial calibration once the SVT is in space, with the segments rarely moving after this. This is because the mirror will not have to overcome gravitational or wind loading effects like its terrestrial counterparts.[4.17]

4.4.2.2.2 – Super-Module and FGS The FGS as discussed in section 4.3 will be placed around the perimeter of the space viewing mirror for reasons discussed in section 8.6.3. This structure is unable to be moved due to the sensitive hardware and the high resolution called for by the SVT. Therefore, a larger module was necessary to house the SVM. The resulting super-module expanded the module length from the original 2 m to 5 m, and increased the volumetric equivalent of one-meter-long modules that it occupies from 15 to 25. The large size of the super-module created the need for a second launch to deliver the SVT in the super-module to the empty space on the core of the optics segment that was delivered in a previous launch. The pointing of the SVT would be controlled by the propulsion system and ACS. Since the mirror assembly will have no self-moving mechanisms, the entire spacecraft much rotate for the mirror to focus on one object as it orbits the Earth for an extended period of time. As Hubble has relatively long exposure times to capture the images of the galaxy, the exposure times are anticipated to be of the same magnitude for the SVT. For most of the objects the SVT will focus on, exposure times will vary between 0.75 to 1.5 hours. This will increase when looking far-field. At one point Hubble focused on a far-field object for nearly two weeks.[4.18] The SVT’s super-module will be on the opposing side of the core from the EVT as depicted in Figure 4.14.


Final Report 40

Figure 4.14: Opposing placement of SVT and EVT

This placement allows the EVT to be pointed towards Earth while the SVT will be looking space-ward. Most of the time, however, these telescopes will not be employed at the same time. The computer system is capable of storing and transferring the data from both systems simultaneously. However, the EVT prevents this. The mission of the EVT is to obtain complete coverage of the Earth. Since the coverage of the EVT is less than the distance between adjacent ground tracks of the satellite, it will constantly be required to maneuver on its three axes to cover the distance between them. These movements will cause torques on the system, causing movements in the system counterproductive to the SVT. The SVT, which has longer exposure times than the EVT, must stay focused on an object for an extended period of time. An extremely precise FGS was chosen to obtain a high accuracy of resolution for the space images. If the system is moved slightly, the SVT will not be able to obtain satisfactory images.

4.4.2.2.3 – CCD Chip Assembly and Structure The CCD chip configuration for the Odyssey imaging hardware is equivalent to Hubble’s in resolution and even surpasses it in viewing area. The four postage-stamp sized chips in the CCD chip assembly aboard the Odyssey each measure 1024 x 1024 pixels. These CCD chips are sensitive to the extremely faint light enabling the imaging of distant galaxies. Light that is one billion times fainter than the naked eye can see can be imaged by this system. This is an upgrade from Hubble’s 800 x 800 pixels. Referring to the calculations found in Appendix A.4, this upgrade increases the plate scale of the SVT, increasing the area of the sky observed. This results in equivalently clear pictures of a larger area. The CCD support structure was designed in a tripod configuration for stability and to support and properly orient the CCD chip assembly. To allow for an easily deployable system that can be folded for launch then extended after the mirror has been docked, a new technology known as shape-memory material was chosen. The material NiTi (Austenite), also known as NiTinol, is the most common shape-memory material employed in the aerospace industry. It is currently being tested on the Boeing QT2D

SVT

EVT


Final Report 41

engine in the form of temperature-actuated chevrons. As the temperature of the outlet gasses increases, the chevrons move to control air mixing.[4.19] There are two forms of NiTinol: martensite, a less-ordered (flexible) form and austenite, a more-ordered (solid) form.[4.20] During launch while packaged in the super-module, the three struts that will comprise the CCD structure will be in the martensite form so they will be able to bend and fit safely atop of the beryllium mirror. While being used as solid structures to hold the CCD chip assembly in place, the struts will be in the austenite form. The “parent shape,” or the ultimate desired shape of the struts would be determined during manufacturing. The material would be heated to 500 °C, held for five minutes, and then rapidly cooled. This causes the atoms to arrange themselves in a pattern that is reverted to when the transition temperature is reached. When the SVT is ready to use, the struts must be heated to approximately 166 °C via wires running through them. This would cause the struts to return to its parent shape.[4.21] Each leg of the tripod was designed as a tube to allow for the structure to carry the desired load while minimizing the mass. The minimum mass for Austenite was obtained using the Genetic Algorithm, or GA, optimization technique.[4.22] More about this method is discussed in Appendix A.4. The design shown in Figure 4.15 was used in the GA code, varying diameter d of the struts and their thickness, t.

Figure 4.15: CCD Support Structure[4.23]

The total mass of the CCD support structure is 12 kg, with a 9.65 mm diameter rod and a thickness of 3.326 mm. This supports a 2.26 kg load that can be applied by the CCD before failure would occur via buckling. This incorporates a safety factor of twice the max anticipated load.

4.4.2.2.4 – Manufacturing For manufacturing of the segmented Beryllium mirror of the SVT, Corning OCA or can be contracted. Corning OCA is able to polish the Beryllium mirror segments to the specific wavelengths of visible light that are to be observed by the Odyssey.[4.24] Axsys Technologies, Inc. is another possible contractor, being the company that is currently manufacturing the


Final Report 42

Beryllium segmented mirror to be placed on the James Web Space Telescope.[4.25] The NiTinol for the struts of the CCD support structure will come from Dynalloy, Inc.[4.26] The CCD chips are easily found through GlobalSpec, since they are currently employed in other aspects of technology such as cameras and recorders. 4.5 – Cost Estimation The estimated costs of the portions of the SVT design are summarized in Table 4.3.

Table 4.3: SVT cost estimate

Beryllium Mirror $13,564,800FGS $30,000,000CCD chips $5,000 NiTi CCD Supports $9,300 Aluminum supports for mirror $130 Total Cost of SVT $43,579,000

A further breakdown of these costs can be found in Appendix A.4.5.

Table 4.4: EVT Cost Breakdown Part Cost

Primary Mirror $7,621,000Secondary Mirror $1,056,000CCD chips $9,500Structure $1,262Total Cost $8,687,762

The main mirror of the primary mirror is design primarily the same way as the Hubble Space Telescope’s, so the cost estimation for the mirror can be used (Equation 4.2). But the mirror composition with new manufacturing techniques the cost is 80% of what it would cost to create a normal Hubble type mirror. So the primary mirror costs $7.6 million and the secondary mirror is $1.1 million.

2/10$ mM (Equation 4.2: Mirror Cost Estimation)[4.5]

The CCDs for the EVT are from a paper that explains how CCDs are incorporated into an orbital telescope.[4.2] The cost of the structural components is based off the manufacturing cost of T7075-T6 Aluminum. Table 4.5 gives a breakdown of the cost of each component, since every structural component in the EVT is composed of the same aluminum. The current cost of aluminum is kg/2.2$ .


Final Report 43

Table 4.5: Cost Breakdown by Structural Component

Optics Quantity Mass (each)

CBE Total (kg) Cost ($)

Bipod Mounts to Optics Bench 3 9.500 28.500 $80Hexapod Focus Mechansim 1 11.000 11.000 $31Hexapod Focus Mechansim (OPT 1 11.000 11.000 $31Subtotal 50.50 $142

Telescope StructureSecondary Mirror Housing & Struts 1 31.700 31.700 $89Optics Bench Flat & Curbs 1 92.000 92.000 $259Primary "Cold Stovepipe" Light Baffle 1 10.590 10.590 $30Optics Lower "Coffin" 1 23.750 23.750 $67Optics Bench to Bus Mounting Strut pairs 3 26.500 79.500 $224

Subtotal 237.54 $670STRAY LIGHT BAFFLE AssemblyAl Baffle Tube and Vane Assembly 1 112.8 141 $398Primary (front) Solar Array 1 15.7 15.7 $44Cooldown (rear) Solar Array 1 3.00 3 $8

SUBTOTAL 159.70 $450

Total Mass 447.74 1262.806

4.6 – Summary – Lessons learned and future systems Hubble has shown to be a valuable resource for technological advances, necessitating the design of a replacement for the coming years when Hubble will no longer be in service. Basing a majority of the starting designs for the SVT on the Hubble while incorporating new technologies is justified; there will be potential customers and an eager audience for such a product. This amplifies the importance of the design of the SVT subsystem, suggesting the assignment of more than one engineer to the design of it on future projects. In hind sight, what limited the Odyssey design team greatly was the physical connection of the SVT to the same segment that housed the EVT. Both these systems are extremely sensitive, power-demanding and require different orientations. If the SVT is to focus on an object for an extended period of time, the EVT will be unable to move during that entire time so as to not disrupt the system. If these two imaging subsystems were unconnected, such as in a fractionated system, this problem would be averted. Another limitation was the lack of ray tracing software, making it nearly impossible to calculate the focal length and other vital metrics for a dual-mirrored system. The ease of these calculations for a single mirrored system allowed for the implementation of it into The Odyssey Project. The availability of such software would have broadened the horizons of the project, allowing for a cost and mass comparison between the two choices. The Odyssey’s SVT employs CCD chips that can look at a limited range of the spectrum, capturing images mostly in the visible light range. The use of a variety imaging hardware


Final Report 44

simultaneously would allow the system to capture wider variety of wavelengths in its images. This suggests the use of infrared detectors and other such hardware that would be interchangeable with the visible light imaging hardware. With the advancement of technology and the ability to see the faintest of light galaxies away, the future for space observation is brighter than ever. The future points toward modern missions, clearly defined and focused towards the objectives of space exploration.


Final Report 45

4.7 – References [4.1] Weisshaar, Terrence. Purdue Advanced Research Projects Affiliates (PARPA) REQUEST

FOR PROPOSAL: Modular, Fractionated Space System Advanced Concept Technology Demonstrator." Purdue Advanced Research Projects Affiliates (PARPA) Mission Opportunity Statement. West Lafayette: Purdue University, 2006.

[4.2] "Assembly Slip Ring." GlobalSpec. 11 Oct. 2006 <http://motion-

controls.globalspec.com/Industrial-Directory/assembly_slip_ring>. [4.3] Larson, Wiley J., and James R. Wertz, eds. Space Mission Analysis and Design. 7th ed. El

Segundo: Microcosm, Inc., 2005. [4.4] Dewandre, Thierry M., Joachim J. Schulte-In-Den-Ba umen, and Emmanuel Sein, comps.

Space Optics 1994: Space Instrumentation and Spacecraft Optics. SPIE: Bellingham, Wash, 1994.

[4.5] Ashby, M F. Engineering Materials: an Introduction to Their Properties and Applications.

Oxford: Pergamon P, 1980. [4.6] MatWeb, Your Source for Materials Information." Matweb: Material Property Data. 3 Dec.

2006 <http://www.matweb.com>. [4.7] Malacara, Daniel, and Brian J. Thompson, eds. Handbook of Optical Engineering. New

York: Marcel Dekker, 2001. [4.8] Born, Max, and Emil Wolf. Principles of Optics; Electromagnetic Theory of Propagation,

Interference and Diffraction of Light. 7th ed. Cambridge UP, 1999. [4.9] "Resolution 101." HubbleSite. NASA. 10 Nov. 2006

<http://hubblesite.org/the_telescope/nuts_.and._bolts/res101.shtml>. [4.10] "Hubble is a Reflecting Telescope." HubbleSite. NASA. 9 Nov. 2006

<http://hubblesite.org/hubble_discoveries/hstexhibit/telescope/about.shtml>. [4.11] "Postcards From the Edge." HubbleSite. NASA. 15 Oct. 2006

<http://hubblesite.org/the_telescope/nuts_.and._bolts/instruments/wfpc2/>. [4.12] Finley, John P. Personal interview. 6 Nov. 2006. [4.13] "The Primary Mirror." Jet Propulsion Labs. NASA. 14 Nov. 2006

<http://pds.jpl.nasa.gov/planets/captions/hubble/hst01.htm>. [4.14] "Hubble's Amazing Optics." HubbleSite. NASA. 30 Nov. 2006

<http://hubblesite.org/the_telescope/nuts_.and._bolts/optics/>.


Final Report 46

[4.15] "James Webb Space Telescope Advanced Mirror Demonstrator Tests Under Way At NASA's Marshall Center." Marshall Space Flight Center. NASA. 15 Nov. 2006 <http://www.nasa.gov/centers/marshall/news/news/releases/2003/03-076.html>.

[4.16] "Plate Scale and Focal Lengths: a Tutorial and Worked Example." John Lucey Astrolab.

Durham University. 1 Dec. 2006 <http://www.dur.ac.uk/john.lucey/astrolab/plate_scale_updating.html>.

[4.17] "JWST." James Webb Space Telescope. Ball Aerospace. 30 Nov. 2006

<http://www.ballaerospace.com/jwst.html>. [4.18] "Hubble Heritage Archive - 2002." Journey Through the Galaxy. 23 Oct. 2006

<http://filer.case.edu/~sjr16/advanced/archive_hst_2002.html>. [4.19] Burnett, Bob. "A Boeing-Led Team is Working to Make Quiet Jetliners Even Quieter."

Boeing Frontiers. Boeing. 15 Nov. 2006 <http://www.boeing.com/news/frontiers/archive/2005/december/ts_sf07.html>.

[4.20] Kubiak, John. Aerospace Applications of Shape Memory Alloys. AAE 454, Fall 2006,

Purdue University. [4.21] Lin, Richard. "Shape Memory Alloys and Their Applications." Stanford University. 20

Nov. 2006 <http://www.stanford.edu/~richlin1/sma/sma.html>. [4.22] Crossley, William et al. genetic.m Fitness function minimization code. 20 July 2004. [4.23] Crossley, William. AAE550 Homework 3 A Combinatorial Problem with the GA. Fall

2006 [4.24] “Beryllium Optics." TechSearch. Corning OCA. 1 Dec. 2006

<http://www.mdatechnology.net/techsearch.asp?articleid=175>. [4.25] "Axsys Technologies Chosen to Produce Beryllium Mirror Substrates for Use on the

James Webb Space Telescope." News. Axsys Technologies, LLC. 1 Dec. 2006 <http://www.axsys.com/pr-030922.php>.

[4.26] “NiTinol Price Guide." Dynalloy, Inc. 15 Nov. 2006

<http://www.dynalloy.com/PriceGuide.html>.


Final Report 47

5.0 – Orbit Selection There are six basic elements that define a spacecraft’s position in a specific orbit. The orbit is defined by the argument of periapsis, eccentricity, semi-major axis, inclination, and right ascension of the ascending node. True anomaly is the sixth element that defines the spacecraft’s position in the orbit 5.1 – Scope and Purpose This section considers the orbit design and final orbit selection of the Odyssey. The purpose of this section is to discuss the evolution of the design from its inception to the final orbit selection. First, the tasks, functions, and requirements of the orbit as well as significant design problems and issues are identified. Then, the trade-offs and design selection process is discussed. To conclude, the final orbit is discussed as well as how it responds to the requirements and design problems. 5.2 – Tasks, Functions, Requirements and Design Methodology In the selection of the final orbit, three main tasks were completed. These included developing the trajectory for the best performance of the mission, determining the velocity changes to insert the satellite into the desired orbit, and determining the velocity changes required for simple plane change maneuvers. The function of the orbit is to provide a trajectory suitable for both Earth and space imaging. While this was the only function of the orbit, it posed a challenging design problem because of the dual imaging requirement. In addition to the system requirements, several important considerations regarding the system’s optical capabilities influenced the selection of the final orbit. These are discussed in sections 5.3.1 and 5.3.2. According to the Mission Opportunity Statement, the Odyssey must operate in low-Earth orbit. The Mission Opportunity Statement also requests that the Odyssey have orbital plane change capabilities. This requirement is discussed in detail in section 5.3.3. Further requirements were added in the systems requirements document. It states that the orbit must be appropriate for both Earth and space imaging, the system must provide complete coverage of the Earth, and the system must maintain an altitude of 785 km. Discussions of the altitude requirement can be found in section 5.3.2 and 5.3.4. The design methodology consisted of six steps:

1) Determine system requirements 2) Identify significant design problems and issues 3) Study similar systems 4) Choose baseline orbit 5) Iterate on baseline orbit given input from other teams 6) Determine final orbit


Final Report 48

5.3 – Orbit Selection - Problems and Issues The Odyssey’s system requirements led to several important considerations that affected the selection of the final orbit. The considerations due to the Earth imaging requirement are discussed in section 5.3.1, and the considerations due to the space imaging requirement are discussed in section 5.3.2.

5.3.1 – Earth Imaging The Odyssey’s Earth imaging requirement was the primary design driver of the satellite’s orbit. The Earth imaging requirement imposed several restrictions on the orbit selection; the space viewing requirement imposed only one. For Earth imaging, the several important considerations influenced the selection of the final orbit. First, for a valid comparison of images taken at a location on different dates, the images must be obtained from the same altitude under similar illumination conditions. To take consistent repeat images of Earth targets, it is also desirable that the orbital path repeat after a reasonable, whole number of days. This can be accomplished with a phased orbit. A phased orbit is an orbit that, by careful selection of the semi-major axis, completes an integer number of revolutions in an integer number of days. Finally, the orbit must provide complete coverage of the Earth. This requirement is discussed in further detail in section 5.3.3. Table 5.1 shows the corresponding orbit characteristics to the stated Earth imaging desires.[5.1]

Table 5.1: Corresponding orbit characteristics to Earth-imaging desires Earth imaging desires Corresponding orbit characteristics Images obtained from constant altitude Circular Consistent illumination conditions Sun-synchronous Repeating orbital path “Phased” Complete Earth coverage Near-polar

5.3.2 – Space Imaging The Mission Opportunity Statement states that the system must operate in Low -Earth Orbit (LEO). The Odyssey requires the same space imaging capability as the Hubble Space Telescope (HST), and both systems operate in LEO. Consequently, HST provides a good reference for the space imaging challenges and opportunities of the Odyssey. The main restriction placed on the orbit design by the Odyssey’s space imaging requirement is that the system operates at a sufficient altitude. A sufficient altitude is defined here as one which places the system above the distortion of the Earth’s atmosphere for unrestricted visibility into space. For space-imaging, a satellite in LEO will have two Continuous-Viewing Zones (CVZ) parallel to its orbital plane in which objects remain unblocked by the Earth for up to seven weeks.[5.2] Figure 5.1 gives a pictorial representation of HST’s CVZs. The Odyssey’s orbit (i ≈ 98.5º) is more highly inclined than HST’s orbit (i ≈ 28.5º). Therefore, the Odyssey’s CVZ will be shifted by approximately 70º from the CVZ shown in Figure 5.1. For objects outside of the CVZ, space imaging depends on how long a target remains unblocked by Earth.[5.3]


Final Report 49

Figure 5.1: HST’s Continuous zone viewing[5.3]

Like HST, it will be necessary to have a 50º Sun-exclusion zone for the SVT to protect it from damage. For this reason, the satellite will not be able to image Mercury and Venus because they fall inside this zone.

5.3.3 – Trade-off: Simple Plane Change Maneuvers An original system requirement of the Odyssey project was that the satellite would have the ability to perform simple plane change maneuvers. The main purpose behind this requirement was increased responsiveness of the satellite and, more importantly, the ability to take images anywhere on the Earth. Complete coverage of the planet is necessary for an Earth imaging satellite; however, plane change maneuvers are expensive. The cost of a modest plane change can be seen in Figure 5.2. In this figure, vi is the velocity of the satellite in the original orbit (vi ≈ 7.47 km/s). The Δi is the desired change in inclination. Changing the vehicle inclination by 10º requires a ΔV of about 1.3 km/s. In other words, a single impulsive maneuver requires a ΔV greater than the total ΔV budgeted for the satellite for an entire year. Refer to section 8 for more information about the ΔV budget. However, with careful planning, the selected orbit can be chosen such that no plane change maneuvers are required for complete Earth coverage.


Final Report 50

Figure 5.2: Delta-v required for simple plane change

If the Field-of-View (FOV) of the optical instrumentation is larger than the maximum width between adjacent ground tracks, a near-polar orbit will provide complete Earth coverage.[5.1] Without plane change capability, the satellite is more cost effective. It will not be necessary to carry the extra fuel aboard to perform expensive plane change maneuvers. The choice of a near-polar orbit is an attractive alternative to the original system requirement. 5.4 – System Orbit Description The selected altitude, 785 km, operates at a higher altitude than HST (hHST = 600 km). This easily surpasses the “sufficient altitude” needed for space imaging. It also has additional desirable characteristics. The atmospheric drag, for example, is very low at this altitude, decreasing the need for orbital maintenance maneuvers. For an extended discussion of atmospheric drag and maintenance maneuvers, see section 8. The altitude is also a good choice for sun-synchronous orbits (usually 600 to 800 km), discussed next.

5.4.1 – Sun-synchronous Orbit (SS-O) The final orbit was chosen to be an SS-O because it allows points on the Earth to be crossed at the same local time. This provides consistent illumination conditions for Earth-imaging. An SS-O takes advantage of the Earth’s equatorial bulge by using the J2 effect to match the precession rate of the satellite’s orbit to the Earth’s rotation around the sun, or approximately 360º/year (0.9856º/day). The inclination required for an SS-O can be determined from Equation 5.1.[5.4]

( )( ) 222/714 11006474.2

cos 2

−− −−

Ω≅

eaxi J

(Equation 5.1)

0 2 4 6 8 100

0.2

0.4

0.6

0.8

1

1.2

1.4

Change in inclination, degrees

| Δ v

|, km

/s

Delta-v required for changing vehicle inclination

Δ v = 2vi sin(Δ i/2)


Final Report 51

In this equation, the variation of the Right Ascension of the Ascending Node (RAAN) due to J2

2JΩ is 0.9856º/day. The semi-major axis of the orbit (a = 7163.19 km), and the eccentricity (e = 0), result in an inclination is 98.54º.

5.4.2 – Orbit Characteristics Based on considerations mentioned above, the final selected orbit is a 25D358R SS-O. This designation indicates the orbit path repeats every 25 days and 358 revolutions, and is sun-synchronous. It is also near-polar, circular, and operates at an altitude of approximately 785 km. It has an equator crossing time of 10:30am. Table 5.2 shows the relevant orbital characteristics.[5.5] Detailed calculations can be found in Appendix A.5.

Table 5.2: Orbital Characteristics Altitude h = 785 km Eccentricity e ≈ 0 Inclination i = 98.54º Mean local time of equator crossing

MLT = 10:30am

Orbital period P = 100.56 minutes Revolutions per day 14.32 revs/day Length of orbital cycle D = 25 days Number of revolutions before orbit repeats

R = 358 revs

Westward drift between successive ground tracks

∆L = 2822.19 km

Maximum width between adjacent ground tracks

∆l = 112.89 km

Rate of change of RAAN 2JΩ = 0.9856º/day

Rate of change of argument of perigee

2Jω = -0.5506º/day

Figures 5.3 and 5.4 show the final orbit of the Odyssey. These were produced using Satellite Tool Kit (STK). The STK model was produced using a J2000 coordinate system and “Earth full” propagator. The model was used to confirm calculations. Refer to Appendix A.5 for more information.


Final Report 52

Figure 5.3: Odyssey orbit

Figure 5.4: Odyssey ground track [1 day]


Final Report 53

5.4.3 – ΔV for Orbit Insertion The launch vehicle, Falcon 9, will launch the Odyssey from Vandenberg Air Force Base directly into its final orbit. The ΔV required for orbit insertion is determined by the final velocity of the satellite in its orbit (VLEO) plus the sum of the ΔV losses due to gravity, drag, and steering.[5.6] At 785 km, the VLEO = 7.47 km/s. The other ΔV’s for orbit insertion given in Table 5.3 are average values.

Table 5.3: Δv required for orbit insertion VLEO 7.47 km/s ΔVgrav 1.39 km/s ΔVdrag 0.12 km/s ΔVsteering 0.15 km/s ΔVTOTAL 9.12 km/s

5.4.4 – Launch Window The satellite will be launched from Vandenberg Air Force Base. The launch must occur when, given the rotation of the Earth, Vandenberg crosses the Odyssey’s orbit plane. It is possible to place the Odyssey in orbit either in the ascending part, by launching northwards, or in the descending part, by launching southward. The southward launch is rejected because of safety considerations.[5.7] The launch of the Odyssey will be in a northerly direction, into the ascending node, when Vandenberg is near the local time of the orbit’s ascending node. The launch must take place within a small launch window lasting 5 minutes to get Odyssey into the correct orbit plane. The target time is at the beginning of the launch window so that if there are delays in the countdown it is still possible to launch the satellite. Outside of the launch window, the launch vehicle’s ΔVsteering would be too high.[5.7] 5.5 – Summary – Lessons Learned and Outlook for the Future Determining the final orbit for the Odyssey allowed for the study of different orbit types that operate in LEO. In selecting the orbit, the design team was able to determine the effect of different Earth and space requirements on the orbital characteristics. Furthermore, the benefits and challenges of SS-Os were studied in detail. Future work on the orbit might include a detailed analysis of the availability of space targets over a period of time.


Final Report 54

5.6 – References [5.1] “Spot on its orbit.” 1 October 2006.

<http://www.spotimage.fr/automne_modules_files/standard/public/p229_fileLINKEDFILE_Spo-on-its-orbit.pdf>.

[5.2] Wells, J. G. “Hardware Emulation and Real-Time Simulation Strategies for the Concurrent

Development of Microsatellite Hardware and Software.” University of Toronto. 2001. 1 December 2006. <http://www.collectionscanada.ca/obj/s4/f2/dsk3/ftp05/MQ62899.pdf>.

[5.3] “Hubble Space Telescope: Servicing Mission 3A Media Reference Guide”. 27 December

1999. 22 October 2006. Prepared by Lockheed Martin. <hubble.nasa.gov/a_pdf/news/SM3A-MediaGuide.pdf>.

[5.4] Larson, Wiley J., and James R. Wertz, eds. Space Mission Analysis and Design. 3rd ed. El

Segundo: Microcosm P, 2005. [5.5] Boain, Ronald J. “A-B-Cs of Sun-Synchronous Orbit Mission Design.” 14th AAS/AIAA

Space Flight Mechanics Conference. 8-12 February 2004. 20 November 2006. <http://trs-new.jpl.nasa.gov/dspace/bitstream/2014/37900/1/04-0263.pdf>.

[5.6] Humble, Ronald W., Gary N. Henry, and Wiley J. Larson. Space Propulsion Analysis and

Design. New York: The McGraw-Hill Companies, Inc., 1995. [5.7] Brel, E. Cabrieres, B. et al. “Spot 4’s orbit.” 06 June 2000. 15 November 2006.

<http://spot4.cnes.fr/spot4_gb/orbite.htm>.


Final Report 55

6.0 – Communications and Data Handling All data transfer related to the satellite mission is controlled by the communication and data handling subsystems. Both the communication and data handling subsystems will be discussed in this section of the report. The communication system consists of hardware onboard the satellite and the links used to transfer data while the data handling subsystem consists of the computers and memory onboard as well as the overall architecture used to transfer data. 6.1 – Scope and Purpose The purpose of the communication and data handling subsystems is to provide a robust system that can handle the demands of satellite operations. The communication and data handling systems define the functional capacity of the spacecraft and how much science can be conducted and transmitted back to Earth. Computers and data storage devices make up the main components of the onboard data handling system. Computers control all autonomous operations of the spacecraft including propulsion system, attitude control, and communication links. The computer systems are also responsible for storage of all data until it can be transmitted to the ground station for storage. Each computer consists of radiation hardened processors and memory in a protective enclosure. They are linked and controlled by software that distributes the processing load and determines where data is stored. A traditional consolidated computer system would involve a primary computer and a backup computer which would only be activated if the primary computer fails. A distributed system uses several less capable systems to achieve the redundancy without carrying a designated backup computer. Each computer in the distributed system shares the workload and in the event of a computer failure the other computers can sustain the spacecraft until a replacement is sent. The communication system is composed of transceivers and antennas which establish a wireless link to a specified target. A transceiver reduces size and mass requirements by combining the transmitter and receiver into a single unit and sharing circuitry and power amplifiers. The antenna transmits and collects electromagnetic (EM) waves and converts them into signals compatible with the computer systems via a transceiver. The details of both subsystems are discussed in the following sections. 6.2 – Tasks, Functions, Requirements and Design Methodology The primary function of the communication and data handling subsystems is to process, store, and transmit all science data to Earth in addition to receiving tracking, telemetry, and command (TT&C) information. The data handling system’s function is to process and store all onboard data while the communication system’s function is to establish the data links.


Final Report 56

The payload team was responsible for ensuring that the communication and data handling systems onboard Odyssey and on the ground meet mission requirements specified in the mission opportunity statement and defined by the team. The top level requirement defined in the mission opportunity statement was that satellite operations must be controlled from West Lafayette, Indiana. To meet this requirement the team added the requirements of continuous communication coverage and a ground station located in West Lafayette. Continuous coverage will be achieved through NASA’s Telecommunications and Data Relay Satellite System (TDRSS). There were no specific data handling requirements established in the mission opportunity statement. All requirements imposed on the data handling system were put in place by the team to meet the needs of individual systems. The imposed requirements pertain to the processor power and memory capacity of the computer systems onboard the satellite. Ground station data handling requirements were also imposed on the storage capacity for archiving all the data obtained from the satellite. The communication system has several tasks it must perform to do its job. The first task for the communication system is to convert the information sent from the computer into a signal that can be transmitted back to the West Lafayette Ground Station (WLGS). The communication system is also responsible for establishing and maintaining S-band links with TDRS for TT&C. For the high speed data download, the communication system must establish and maintain Ku-band links with TDRS. The main task of the data handling system is to control all spacecraft components including the communication system and data links with TDRS. The computers are also required to process and store all imaging data collected by both telescopes. During docking procedures the computers and communication system will control the approach of the module or segment. Sufficient storage capacity must be provided for data until it can be downloaded to WLGS. In between communication links the operational and imaging data needs to be stored onboard Odyssey. The computers will distribute the data to the memory systems where it will be stored until WLGS confirms the data is stored on the ground. The WLGS is the command center and data storage facility for the Odyssey mission. All commands that need to be sent to the satellite will be sent from WLGS via TDRSS to Odyssey and data will be sent via TDRSS back to WLGS for analysis and storage. A storage capacity of approximately 200 TB per year of operation will be required on site to save the raw scientific data collected by Odyssey. By compressing the data using standard JPEG format the data can be compressed to approximately ¼ the size of the raw data. The total storage requirements for raw and compressed data over 10 years are shown in Figure 6.1.


Final Report 57

1 2 3 4 5 6 7 8 9 100

200

400

600

800

1000

1200

1400

1600

1800

2000

Years in Operation

Imag

e D

ata

Dow

nloa

ded

(TB

)

WLGS Data Storage Requirements

Uncompressed DataCompressed Data (JPEG)

Figure 6.1: WLGS storage requirements

The design of the communication system started with the requirement to provide continuous communication coverage of the satellite’s orbit. This requirement led to the use of TDRSS to direct all communications to WLGS. The satellite communication frequencies and data rates were selected to be compatible with TDRSS. The communications link equations were used for several communication scenarios to determine an appropriate antenna size. The other hardware components were then chosen based on the link requirements and available technologies. The system’s computing and data storage needs were used to initially size the data handling system. The large amount of image processing led to an investigation of state of the art radiation hardened computer components. Data handling systems currently in operation were used as a benchmark to determine appropriate computer power and memory capacity. Odyssey’s computer processors and memory were then selected based on current state of the art technology and availability. 6.3 – Choices The first choice that was made in the design of the communication system is whether Odyssey will communicate directly with West Lafayette or through a relay system. A relay system will allow Odyssey to communicate with West Lafayette even when it is not in orbit over Indiana. The type and size of the antenna was chosen next followed by the link parameters such as data rate, frequency and power. If the result is a link margin less than 3 dB then the antenna and link parameters must be changed until a satisfactory margin is achieved. The typical properties of the antennas such as peak gain were also examined for feasibility.


Final Report 58

The type of computer system and capabilities were the primary choices made in the design of the data handling system. The two methods considered for the type of computer system were centralized and distributed systems. In a centralized system there is a single computer that handles all data processing without additional computers. A distributed system has two or more computers which share the processing load and have no single point failure. The power and number of processors as well as the amount and type of storage was chosen in the design of the data handling system. 6.4 – Design Evolution The design of the communication system began with a concept of system that communicated directly with an antenna in West Lafayette. Once it was decided that more frequent access would be required; relay systems were investigated. NASA’s TDRSS was selected as the best option so the communication system was designed to be compatible with first and second generation Tracking and Data Satellites (TDRS). A parabolic reflector was selected as the antenna type based on the frequencies and gain required for the communication link. A second antenna was added on the opposite side of the satellite to enable communication in any orientation. The capability for Ku-band communications was added for high data rate capability to handle the transmission of imaging data. The data handling system design began with a debate on whether to use a centralized or distributed computing system. Once a distributed computer system was selected, the computer and memory systems were sized based on the systems onboard the Hubble Space Telescope (HST). It was determined that the Odyssey would need more capable systems due to the dual telescope design. The computer system components were sized based on information on current state of the art components. To improve the reliability, additional computers were added to the system for redundancy and increased capability. 6.5 – Telecommunication Design The telecommunication design includes selecting the size and type of antenna and specifying hardware details. The capabilities of the hardware required to meet the mission requirements are discussed in the following sections.

6.5.1 – Design Approach The communication system was designed to be robust and limit the restrictions placed on the system. The goal was a system that enhances the spacecraft’s capabilities and can handle expansion of the satellite. The system was designed to handle Tracking Telemetry and Control (TT&C) links and data download links. The two links have different requirements and require components with different capabilities. The design task is to reduce mass by incorporating new technologies and using multifunctional components.


Final Report 59

6.5.2 – Requirements – Trade Studies To meet the requirement of continuous communication coverage a relay system is needed to control the satellite from WLGS. The options considered were a private global ground station network and NASA’s TDRSS. TDRSS was selected for its data transfer capabilities and reliable access. NASA keeps an on orbit spare incase a satellite fails and communicating via TDRS in space limits the threat of weather disrupting the links. A simulation was run in Satellite Tool Kit (STK) which confirmed TDRSS meets the constant communication coverage mission requirement (Figure 6.2).

Figure 6.2: STK link simulation screenshot

TRDSS offers a wide range of frequencies from S-band to Ka-band for Odyssey to potentially use. The multiple-access S-band antenna offers reliable access at relatively low data rates up to 10kbps and up to 300 kbps via the single access S-band antenna.[6.1] These links work well for the tracking, telemetry, and control (TT&C) links which are frequent and don’t involve large amounts of data transfer. The mass penalty for adding a high speed Ku-band capability was compared to the benefits of a high data rate link. By using the same antenna for both Ku and S-band links the only change to the system is addition of Ku-band transceivers. Downloading images and other engineering data in less time enables the system to spend more time imaging which led to the decision that the mass penalty is worth the capability. The antenna size was analyzed using a Matlab code shown in appendix A.6.1.4 that plots the antenna size verses the link margin. The results were used to select the smallest antenna that would provide a margin of at least 3 dB for all four link scenarios.

6.5.3 – Final Concept – Component Descriptions The communication system provides reliable on-demand access and high speed data transfers via TDRSS. The system consists of two complete communication systems which offer redundancy in the event of component failure and orientation-independent access to TDRSS. Each system consists of a parabolic reflector antenna and two transceivers.


Final Report 60

The antenna is a 0.5 m mesh deployable parabolic reflector built by Harris Corporation. Harris was selected because of their experience with deployable reflectors (Figure 6.3). The antenna is modeled after the single access antenna used by TDRS which is capable of frequencies from S-band to Ka-band. The antenna surface is constructed out of the gold-plated molybdenum wire mesh used by Harris on the TDRS. This construction has been proven to provide an antenna capable of S and Ku band communication. The antennas are mounted on the end of extendable booms in the communications modules and have 180° pitch freedom and 360° rotational freedom. This range of motion allows them to track the TDRS satellites without changing the orientation of the spacecraft.

Figure 6.3: Gold-clad molybdenum mesh parabolic reflector built by Harris[6.2]

The S-band links offer low-speed reliable access via the multi-access (MA) or S-band single access (SSA) antennas onboard TDRS. The S-band uplink operates at a frequency of 2.1064 GHz for the MA antenna and from 2.025 GHz to 2.120 GHz for the SSA antenna. The S-band downlink operates at a frequency of 2.2875GHz for the MA antenna and from 2.200 GHz to 2.300 GHz for the SA antenna.[6.1] The S-band links will use the antenna onboard Odyssey that has a clear view to TDRS. There is one S-band transceiver for each antenna which will handle both transmission and receiving tasks for its antenna. The Ku-band links offer high speed data downloads via the single access (KSA) antenna onboard TDRS. Ku uplinks operate at a frequency of 13.775Ghz and downlinks operate at 15.003GHz via the KSA.[6.1] The Ku-band links will use the same 0.5m antennas onboard Odyssey as the S-band links which reduces the hardware requirements and saves mass. A Ku-band transceiver will be connected to each antenna allowing the spacecraft to send the data to TRDSS independent of orientation.


Final Report 61

6.6 – Data Handling Design The data handling subsystem controls how data onboard the spacecraft is controlled, processed, and stored. The data handling system is comprised of a computer, which also contains the processor as well as the random access and cache memories, and storage memory.

6.6.1 – Design Approach The design approach used for the creation of the computer system was based on improving old technology. The computers onboard Hubble were used as a model and improved to meet the needs of Odyssey. Creating a space-capable computing system requires hardening all of the components against the radiation by adding additional transistors.[6.3]

Hubble computers are currently based off a 150 MHz processor and use 12 GB of memory to perform its mission.[6.4,6..5] Since Hubble, radiation hardening and processing power have improved. These advancements enable the use of more advance processors on Odyssey.

6.6.2 – Final Concept – Components and Specifications Aitech Defense Systems, which produces radiation shielded IBM PowerPC processors, will be contracted to build four radiation shielded computers based on the 2.2 GHz IBM PowerPC 970FX processor.[6.6,6.7] These computers include critical components such as the processor, random access memory, and cache memory. These processors will be housed in two separate computer modules so that each computer module will be easier to replace. Each computer module will hold two enclosures which are capable of housing eight cards such as a computer or 1 GB of storage memory.[6.8, 6.9] For Odyssey, each enclosure will house one computer and seven GB of storage memory for a total of 28 GB. The two computing modules will be separate, decentralized, and will share the processing of data in a distributed layout. Having four 2.2 GHz computers onboard allows for double redundancy, so that if one or two computers fail, Odyssey will still have full mission capability. If a memory card fails, Odyssey would still maintain full mission capability because it carries four extra memory cards. This would not have a severe impact on the mission unless multiple cards were lost at once. Aitech’s computers have a watchdog timer and boot software built in to help detect when executing programs fail and to reset the computer in case of such a fault. The upgradeable boot software contains the boot up information necessary to power on the computers as well as debugging and diagnostic tools. Aitech computers use WindRiver’s VxWorks version 5.5.1 for an operating system.[6.10] The specifications of power, weight, volume, and cost for each full enclosure as well as for the entire system are presented in Table 6.1.


Final Report 62

Table 6.1: Computer System Specifications[6.10]

Per full enclosure System totalPower (W) 47.5 190 Weight (kg) 12 48 Volume (m3) 0.00725 0.0290

Cost $30,000 $120,000 6.7 – Material Cost Estimation Data for the material cost estimation of the communication system is proprietary in nature and estimations were required. The cost estimation is based on the material of the antenna as well as the approximate price of transceivers. Molybdenum mesh is approximately $2150 per square meter.[6.11] For a 0.5 m parabolic reflector the cost for the surface will be about $2000 including the gold for plating the wire surface. Estimates on the cost of the transceivers include development and other costs associated with construction. NASA’s Low Cost Transceiver (LCT2) is being designed for TDRSS S-band access with the goal of lowering the price per transceiver from $250,000 to $75,000.[6.12] Using this cost as a baseline for the estimate. Development of two different transceivers is estimated at $120,000 with an additional $30,000 for the additional units. This cost could be reduced significantly if a satisfactory product becomes commercially available. The estimated cost for adapting an off the shelf transceiver is $28,000. A COTS transceiver could be modified for about $5,500.[6.13] The cost of $8,000 per transceiver is used as an approximation. The total cost of the computing system takes into account the cost of all the computers, storage memory, and enclosures. The cost of each full enclosure is $30,000 bringing the total cost of the computing system to $120,000.[6.10] This cost includes material, labor, and manufacturing costs because computers built by Aitech come pre-assembled. 6.8 – Summary – Lessons Learned and Outlook for Future Designing a communication system requires compromises in size, mass, and ability. The Odyssey communication system was designed with an emphasis on reliability and capability. Combining the dual frequency band capability with the services offered by TDRSS provides Odyssey with both reliable on-demand access and high speed data transmission. One lesson learned in the development of the communication system is that the proprietary nature of the communications hardware made specific details difficult to find. Despite the vast simulation tools available today, experimentation is necessary to determine exact performance of antenna designs. Communication technology continues to advance at an incredible pace. As the hardware advances the data links can carry more data at higher speeds. With the increase in capabilities new opportunities to generate revenue and share data also grow. The modular nature of Odyssey allows the satellite to evolve with the communications technologies with new capabilities being added as needed.


Final Report 63

One lesson learned was that not just any Earth-based computer can be launched into space. A computer for space application must be shielded from the harmful radiation environment found in space. This is accomplished by using more robust transistors that are not as susceptible to radiation.

Another lesson learned was that a distributed computing system is more favorable for use in Odyssey because a centralized computing system would not allow for easy expansion or replacement of the computing system. A distributed system also helps disperse heat and allows for redundant computers to be used. A distributed computing system helps to ensure the survival of Odyssey because should a computer fail, the redundant computers would take over data processing until the failed computer could be replaced. The communication and data handling systems designed for Odyssey meet the specified requirements and provide robust capabilities. Both systems enhance the system and support future expansions to Odyssey.


Final Report 64

6.9 – References [6.1] Larson, Wiley J., and James R. Wertz, eds. Space Mission Analysis and Design. 3rd ed. El

Segundo: Microcosm P, 2005 [6.2] "Commercial Deployable Antenna Reflectors." Harris Government Communication

Systems. 20 Nov. 2006 <http://www.govcomm.harris.com/>. [6.3] “Radiation Resistant Computers.” National Aeronautics and Space Administration. 18

Nov. 2005. <http://science.nasa.gov/headlines/y2005/18nov_eaftc.htm>. [6.4] “Hubble Space Telescope Servicing Mission 3A Media Reference Guide.” Lockheed

Martin. 27 Dec. 1999. National Aeronautics and Space Administration. 27 Dec. 1999 <http://sm3a.gsfc.nasa.gov/media_guide.html>.

[6.5] “Hubble Facts.” Goddard Space Flight Center. 15 June 1999. National Aeronautics and

Space Administration. 15 June 1999. <http://hubble.nasa.gov/a_pdf/news/facts/FS15.pdf>.

[6.6] Aitech Defense Systems Rugged Computer Systems. 2 Dec. 2006.

<http://www.rugged.com/>. [6.7] “IBM PowerPC®970FX RISC Microprocessor Data Sheet.” IBM Systems and Technology

Group. 4 June 2006. <http://www- 306.ibm.com/chips/techlib/techlib.nsf/techdocs/1DE505664D202D2987256D9C006B90A5>.

[6.8] “E900.” Aitech Defense Systems. 2 Dec. 2006.

<http://www.aitechlibrary.com/aitech/protected/e900.pdf>. [6.9] “S990.” Aitech Defense Systems. 2 Dec. 2006.

<http://www.aitechlibrary.com/aitech/protected/s990.pdf>. [6.10] “S950.” Aitech Defense Systems. 2 Dec. 2006.

<http://www.aitechlibrary.com/aitech/protected/s950.pdf>. [6.11] "Product List." Cleveland Wire Cloth. 14 Nov. 2006

<http://www.wirecloth.com/molybdenum.htm>. [6.12] Bull, Barton. Range and Spaceport Science and Technology Workshop, 10 Jan. 2005,

NASA. 23 Nov. 2006 <http://advrangetech.ksc.nasa.gov/Media/10_AFSS_&_LCTT.pdf>.

[6.13] Lu, Richard A. Modifying Off-the-Shelf, Low Cost, Terrestrial Transceivers For. 1 June

1996. Stanford University. 23 Nov. 2006 <http://ssdl.stanford.edu/ssdl/images/stories/papers/1996/ssdl9605.pdf>.

http://hubble.nasa.gov/a_pdf/news/facts/FS15.pdf

http://www.rugged.com/

http://www.aitechlibrary.com/aitech/protected/s990.pdf


Final Report 65

7.0 – Structures and Bus Design The structures subsystem includes the design of the configuration of the spacecraft, the design of the structure, the selection of the materials, and the design of the mechanisms. Discussions on these activities are provided in the following sections 7.1 – Scope and Purpose The purpose of the structure subsystems is to provide the final configuration of the spacecraft. This includes the bus design. In the early phases of the design process, efforts concentrated on the general arrangement of the spacecraft. After defining the final concept of the system, as described in Section 3.0, both the structural support concept and the mechanisms required all flow from the general arrangement provided by the final selected concept. The scope of the structures subsystem covers three areas: the design of the configuration and structure, the selection of materials and the design of spacecraft mechanisms. Mechanisms will be required for deployment from launch configuration to a mission configuration and for other operation purposes. 7.2 – Tasks, Functions, Requirements and Design Methodology The main function is to provide structure to protect the spacecraft's components from dynamic environments during ground operations, launch, deployment, and mission operations. Other functions include providing mounting for all equipment, providing mechanisms for articulation, and controlling the general spacecraft arrangement. The requirements for Odyssey's spacecraft structures include values for strength, structural life, and mass properties derived for the mission and capable of surviving expected environments. The primary structure is defined as the major load path between the spacecraft's components and the launch vehicle.[7.1] In the Odyssey spacecraft the primary structure is the core. The core’s structure is intended to take the expected loads from launch. The anticipated types of loading during launch are axial and lateral acceleration loads and the compressive, bending, and torsion structural loads. Flowing from the initial system concept sketch, the structural support concept was designed to be consisting of the main core with its current octagonal shape. As the design progressed, changes in this concept of support became less global and the general configuration of the system was decided. Secondary structures are defined as structures including support members like the stringers used in the core and the Z-channels used in the propulsion segments to support tanks. Solar panels are also considered secondary structures.


Final Report 66

Requirements for Odyssey's spacecraft materials were to survive ground, launch, and on-orbit environments. These environments impose time-varying applied forces, pressure, humidity, radiation, contamination, thermal cycling, and atomic particles.[7.1] Materials must also help control temperatures. See Section 12, Thermal Control. Competing properties in the material selection process were cost, weight, machinability, and mechanical properties; e.g., stiffness, strength, toughness, fatigue properties, and resistance to corrosion or radiation-erosion. Details about the material selection process will be provided in section 7.5.2.3. Requirements for mechanisms are to provide release and recapture capabilities for the deployable modules and segments of Odyssey. A mechanism is defined as a moving mechanical assembly, such as a hinge, that may be powered electrically or by other means, such as explosives, springs, or other simple devices.[7.1] An essential requirement was to locate the mechanisms in key structural locations were strength and stiffness are sufficient to with stand expected docking or un-docking loads. Another requirement for mechanisms is to operate after lengthy ground testing, which is necessary to show that they will function before and after (and often during) exposure to simulated mission environments.[7.1] Fortunately, the state of the art technology in this area has advanced to the point that some release devices, which are part of the release mechanisms, are now more reliable than the electrical systems that provide the actuation signals.[7.2] There are two types of docking mechanisms used in Odyssey; docking mechanism between modules and docking mechanism between segments. The requirement for Odyssey to be a modular system inherited the need of having these two types of mechanisms and docking schemes. Details are discussed in section 7.7. 7.3 – Design Choices and Drivers The development of the system structure resulted from the fulfillment of the low level system requirements. The requirements affecting the result of the structural configuration and mechanism designs were the system must be: modular, reconfigurable, upgradeable and perform docking operations. The modularity requirement has the greatest effect on the appearance of the satellite in comparison to its effects on cost, performance, and timeline. Modularity forced the components to be housed in separate packages; which at a later date can be removed and replaced for the purpose of upgrading. However, modularity created several obstacles. It resulted in the use of increased material to build the satellite. If all of the components were packaged into a single body covered with a skin it would reduce the material used for the module casings and the skin on the core. Modularity also gave the satellite its appearance; of several modules located radially around the octagonal segments, which are also modular. The configuration also fulfilled an additional requirement of reconfigurability and upgradeability since all components are easy to access. Thus modularity produced the most influential design driver in the structures subsystem.


Final Report 67

Following from the modularity and upgradeable requirements was the docking requirement. The inclusion of docking resulted in the necessary developments of a conceptual docking procedure, described in the ACS section 8.7, and a design concept for docking mechanisms containing the needed interfaces for operation. 7.4 – Design Evolution The design drivers discussed in section 7.3 lead to the initial structure concept. The docking and modularity requirements strongly influenced the first concept developed. The initial designs were the result of the conceptual designs, reflecting their individual unique features. However, with the selection of the final concept the design matured to include possible solutions to the design problems and obstacles, such as the design of a structure to support docking of modules and survive launch in component form. The coupling of the concept selected and requirements of the structures subsystem resulted in the development of three structures and several specialized mechanisms. 7.5 – Segment Configuration The configuration of Odyssey’s segment reflects the system requirement of modularity for the system. Each segment is composed of 1 m modules located on an octagonal core. The modules contain the components for Odyssey’s subsystems; while the core serves as a connection point for each module transferring power, data, and fuel throughout the system.

7.5.1 – Design Choices and Decisions The Odyssey system was required to survive launch, deploy on orbit, dock segments together, and support independent modules. During the design of the segment structure, decisions had to be made as to how to best meet these requirements for a modular system. One such decision was to provide thicker skin than would normally surround a satellite, to serve as shielding against the docking impact, which might hit any position. The modules had to be similarly reinforced, since they were required to be capable of launching on their own and maneuvering in for docking. Another decision made during the design was to provide alignment fins to correct for angular errors on the approach of a docking module. Another decision was whether or not to have the core take the compressive loading, or to have the compressive load transfer along a disposable outer truss. The disposable outer truss would have saved mass on orbit, freeing propulsion efficiency. However, it would have to wrap around the modules, would be wider than the payload fairing adapter, requiring the load to be turned inwards, and would not support the spacecraft once on orbit. The decision to put the structure inside core was made. Our segment developed some distinct characteristics from traditional monolithic satellites due to the constraints of modularity.

7.5.2 – Core Design – Approach, Methodology and Evolution The core segments are the backbone of the spacecraft. The core segments transmit power, computer communications, and fuel along the length of the vehicle. One segment core docks to


Final Report 68

the next segment core as the satellite assembles. All spacecraft modules are docked to the core. In the capacity of primary structure, the core was designed to take the compressive, bending, and torsion loads generated by launch accelerations, with each module hanging off of the core in tension. The heavy compressive members are centered in the core, and the modules can have lighter tension members to hold them to the docking port. This is also advantageous for interfacing with the launch vehicle payload adapter, the plate that the satellite sits on. The compressive load path of the core is within the payload adapter’s radius. All members of the core are composed of titanium, as indicated in our material trade study. The core’s primary structure is intended to take the axial and lateral acceleration and the compressive, bending, and torsion structural loads during launch. Our first model for the primary core structure consisted of two rings of beams and eight stringers between them. The beams take the load of the structure above it, and the stringers anchor the docking mechanism and transfer the loads down the core to the next segment or launch adapter. As the design developed, guidance fins were added, to help align modules during docking. The beam rings became endplates to support the equipment and docking mechanisms inside the core. The stringer cross sections evolved from square box-columns to 45 degree parallelogram box-columns to fit against the sides of the core. Buckling braces were added to the optical segment to increase buckling resistance. Finally, torsion bars were added to handle torsion loading on the structure, and the endplates had holes cut in them to reduce mass. The two previous concepts are shown in Figures 7.1 and 7.2.

Figure 7.1 – Initial Structure Concept Figure 7.2 – Second Concept

7.5.2.1 – Internal Structure and Loads The core structure is composed of stringers, buckling braces, torsion bars, skin, and end-plates. In Figure 7.3 the core structure with guidance fins is shown. The structure was designed to withstand a 6g axial and a 1.5g lateral launch accelerations. Stringers take compression loads


Final Report 69

from the axial acceleration acting on docked modules Compression and tension from the bending forces act on the core structure. Their cross sections were sized to resist yielding from axial forces and buckling from compression. Buckling braces were spaced every meter along the five meter optical segment core to increase buckling resistance. Figure 7.4 illustrates the loading and failure modes considered during analysis.

Figure 7.3: Optical segment core structure

Figure 7.4: Considered failure modes and loading


Final Report 70

The endplates were sized to sit one segment on the next. They hold up the weight of mechanisms internal to the core and transfer the loads from the docking mechanism out to the stringers. Holes were cut in the originally octagonal endplates to reduce the mass. The end plate is shown in Figure 7.5.

Figure 7.5: Endplate

The skin thickness was sized to handle impacts from docking. A center face impact load was assumed as a worst case load. The modules were assumed to have arbitrarily directed lateral accelerations acting on them. Lateral accelerations tangent to the core produced the assumed torsion load. Torsion members were sized to take this load and are located in each face of the core.

7.5.2.2 – FEA Model

Figure 7.6: Two finite element analysis cases

For certain parts of the structure, finite element models were created in ABAQUS, a commercial finite element analysis package, to investigate how the member would react to loading. Samples of two analysis cases are shown in Figure 7.6. Linear elasticity was assumed, which is valid for small deformations of the structure and stresses less than the yield strength of the material.[7.3] The stress and deflection is proportional to the loading under linear elasticity and inversely to a


Final Report 71

stiffness matrix, which is composed of combinations of the dimensions and material stiffness. This finite element model was used to create a sub-model on the systems spreadsheet that could react to changes at the systems level, and dynamically size these members. By varying the load, thickness, diameter and the material Young’s modulus, linear models were constructed for end-plates, skin and alignment fins. For the module and core endplates, equations 7.1 and 7.2 were used to model the maximum stress and deflection:

2

2

tPDCσσ =

EtPDC3

4δδ = Equations 7.1 and 7.2

For the skin and alignment fins, equations 7.3 and 7.4 were used:

2tFCσσ =

EtFC

3δδ = Equations 7.3 and 7.4

The plate constants used for each situation were derived using the finite element analysis model:

Table 7.1: Derived plate constants Situation Cσ Cδ Alignment Fin 4.586 1.03E-12 Skin 1.83 4.53 Trapezoidal Module End-Plate 2.973 0.0472 Final Core End-Plate 0.70 0.0041

7.5.2.3 – Material Choices One important trade study that was conducted was to determine the best material to make each structural member from. Choosing materials to minimize mass is highly desirable, since it can cost thousands of dollars to send a pound of material into orbit. Low thermal expansion coefficients are also desirable, since they minimize stress in the material due to non-uniform temperature. Low thermal conductivities were requested by the thermal team as well, so that heat generated from heat sources within the spacecraft and absorbed by the sun-side will not be propagated throughout the structure. Seven materials were studied and evaluated in our trade study. A summary of their properties is shown in Table 7.2[7.3, 7.4, 7.5]:


Final Report 72

Table 7.1: Summary of material properties Tensile Compressive Young's Poisson's Thermal Thermal Density Yield Yield Modulus Ratio Conductivity Expansion

kg/m^3 Mpa Mpa Gpa W/m-K microm/m-C Aluminum (2219-T851 plate) 2850 414 320 72 0.33 130 23.6 Steel (17-4PH H1150z bar) 7860 634 660 201 0.29 15 10.8

Magnesium (AZ31B H24 sheet) 1170 150 165 45 0.35 96 26 Titanium (Ti-6Al-4V annealed

plate) 4430 828 855 110 0.342 6.6 9 Thornel graphite cloth

composite 1400 103 ~ 13 ** Aluminum-graphite MMC 2500 76.9 109 89 0 190 7.5

Pyrolytic graphite (GE advanced ceramics) 2180 80 100 20 0 300 0.5

The material with the highest strength to weight ratio will provide the lightest tensile rod. Different combinations of material properties are needed to provide for the lightest beams or columns or plates. Their material properties interact differently for different structural roles. These combinations were derived from simple beam, and column buckling equations, and mass calculations. They are given below:

Table 7.2: Minimize these parameters to minimize mass for each type of member Endplates Compressive Rod Tensile Rod Columns Beams

Fyρ

ρ/1

Fcy*

ρ/1

Fty*

Eρ * 3/2Fy

ρ

Fy – yield stress, Fcy – compressive yield, Fty – tensile yield. E – Young’s modulus. Rho – density. *From Peery, the others derived independently[7.6]. The materials were ranked on their ability to take each type of load. They were also ranked for low thermal conductivity and expansion coefficients. The thermal ranking, taken with the structural ranking for each type was scored by a sum (6-ranking)2 rule, and tabulated. Titanium scored the best overall, in each capacity, and was selected as the structural material for all roles within the modules and core. Refer to Table 7.4 for details.

Table 7.3: Scoring of materials Score = (6-ranking)^2*weight Plate Beam Rod Column

Aluminum (2219-T851 Plate) 14 14 14 14 Steel (17-4PH H1150z Bar) 20 24 24 20

Magnesium (AZ31B H24 Sheet) 34 34 25 34 Titanium (Ti-6Al-4V Annealed Plate) 50 50 59 38

Aluminum-Graphite MMC 21 18 17 33 Pyrolytic Graphite (GE Advanced Ceramics) 26 25 26 26


Final Report 73

7.5.3 – Module Design The modules will be launched individually for upgrades to Odyssey and will require an internal structure to support the subsystem components through the launch process. This section will describe the structural analysis required for the modular design.

7.5.3.1 – Internal Structure The modules are designed to hang off of the core, which provides the compressive support for the vehicle. The module members are designed to take tension loads, which enabled them to be lighter than the compressive core stringers. The module structure is composed of a frame, an end-plate on which the mass of the module sits and a tension member connecting the far end of the end-plate to the docking mechanism. The docking mechanism transfers all loads from the module to the core during launch. All members of the module are composed of titanium.

7.5.3.2 – Loads The loading on the module is modeled as a constant pressure on the trapezoidal end-plate, the maximum expected mass inside a module times the axial launch acceleration divided by the area. The loading on the tension bars and frame members is derived from static equilibrium of the end-plate. The rest of the frame is sized similarly to the two frame bars running along the docking face, which takes some of the tension load. The structure of the module is shown in Figure 7.7.

Figure 7.7: Module structure

The structural masses for the components of the spacecraft were computed from the sized structure, and are in Table 7.5. The total structural mass for each launch is tabulated, along with the structural mass for the optical segment core, the 1m optical segment modules, the experiment segment core, the 2m long experiment segment modules, and the telescope super-module. In our spacecraft, the structural mass is 50% of the total spacecraft mass, unlike typical monolithic space systems, which usually have around 21.7%.[7.4] The increase is due to the requirement for


Final Report 74

modularity. Our spacecraft had to be composed of self-sufficient pieces which dock and undock with the core, and the associated docking hardware requires extra mass and structure.

Table 7.4: Summary of structural mass Structure mass per launch

Launch 1 2204 kg Launch 2 2060 kg Launch 3 1720 kg Total Structural Mass 5984 kg Primary structure mass per component Optical Segment Core Structure 308 kg Mechanisms and Cables 730 kg 1m Module Structure x15 1165 kg Super Module Structure 2060 kg Experiment Segment Core Structure 251 kg Mechanisms and Cables 626 kg

7.6 – Structural Design of the Propulsion Segment The objective of this design activity is to design the configuration and the structure of the propulsion segment subject to volume and weight constraints. The structure of the propulsion segment should sustain loads during handling, ground test, and launch.

7.6.1 – Configuration Design The first step in designing the configuration of the propulsion segment was to list all the components that will be in this segment. The configuration design should be made so that mass distribution is made as symmetrical as possible about two perpendicular axes. The propellant tanks should be located near the spacecraft’s center of mass. The next step involves listing all the components. The major components are: three propellant tanks, three pressurizing tanks, and one electric propulsion engine. See section 10 for details about the propulsion system design. To choose the configuration that will occupy the minimum volume, the shape of the tanks has to be chosen. The two available options are cylindrical and spherical tanks. Comparison between spherical tanks and cylindrical tanks of equal volume is made, with a variety of sizes for the cylindrical tank taken into consideration. This comparison showed that spherical tanks are better because they occupy less volume than cylindrical tanks (See Appendix A.7 for details). The limiting constraint on volume was to have a maximum diameter of 3 meters and a maximum height of 2.5 meters. This constraint is determined by the size of the selected launch vehicle fairing (See section 9 on Launch Vehicle System). To support the propellant and pressurant tanks, I-beams are connected to both ends of the surrounding shell of the segment. See Figure 7.8. The tanks were mounted to the I-beams using


Final Report 75

the equatorial mount method where a flange extends around the tank equator. Bolts are used to attach the tanks to the I-beams using support structure of Z-shape.[7.1] See Figure 7.9.

Figure 7.8 - Structural configuration of the propulsion segment.

Figure 7.9 - Mounting method for the spherical tanks.

7.6.2 – Structural Design The purpose of the structure is to support the spacecraft during its life cycle while it is subjected to different loads and operating environments. When in lunch configuration, Odyssey’s propulsion segment is placed at the bottom of the spacecraft. It will therefore be subject to the weight of the entire spacecraft above it while on ground. During launch to orbit, the propulsion


Final Report 76

segment will experience a load of 7g times the mass of the payload above it. See section 9 on Launch Vehicle System. The analysis of the structure subsystem of the propulsion segment is detailed in Appendix 7. The analysis involves choosing the material to be used and the structural geometry. The material that was chosen is Aluminum 7075-T6. Properties of this material are summarized in Table 7.6. Aluminum alloys are used dominantly in the aerospace industry because of their good mechanical properties and low weight. Compared to Aluminum 2024-T3, the 7075 alloys have higher strength than 2024 but lower fracture toughness.[7.7] Being subject to compressive stresses where fatigue is less of a problem, the material for the components of the structural segment was selected to be Aluminum 7075-T6.

Table 7.6. - Mechanical properties of Aluminum alloys at room temperature.

Properties

Material E, (GPa) ν Tensile ultimate stress, (MPa)

Tensile yield stress, (MPa)

Density, (g/cm3)

Aluminum 2024-T3 72 0.33 449 324 2.78 Aluminum 7075-T6 71 0.33 538 490 2.78

The analysis also involves choosing the structural geometry. Wide-flange I-beams and rods with circular cross-sections are used. Wide-flange I-beams are used because of their capability to carry bending moments and efficient distribution of mass. In wide-flange I-beams, material in a beam is located as far as possible from the neutral axis where bending is at a maximum. The vertical web carries most of the transverse shear, with the maximum being at the neutral axis. For sizing the I-beams, the rational method of design that was used is the design on the basis of strength.[7.8] The maximum shear and moment in the each of the beams were determined from the shear and moment diagrams. The maximum bending stress and the maximum shear stress are then used to size the members. See Equations 7.5 and 7.6.

MyI

σ = Equation 7.5

VQIb

τ = Equation 7.6

In Equation 7.5, σ is the bending stress, M is internal moment, y is the perpendicular distance from the neutral axis to the point where the stress is to be calculated, and I is the moment of inertia of the cross-sectional area computed about the neutral axis. In Equation 7.6, τ is the shear stress, V is the internal resultant shear stress, Q =

'

' ' 'A

ydA y A=∫ , where 'A is the top (or

bottom) area of the member’s cross-sectional area defined from the section where b is measured, 'y is the distance to the centroid of 'A .

An iterative process was used to find the ratio (y/I) in Equation 7.5, where in each step the bending and shear stress are checked to be less than ultimate values. The final dimensions are


Final Report 77

shown in Figure 7.10. The width of the top and lower flanges was chosen so that both the propellant tanks (D = 0.96 m) and pressurant tanks (D = 0.33 m) can be supported adequately.

Figure 7.10 - Wide-flange I-beam cross-sectional dimensions.

Rods with circular cross-sections are sized to carry loads without buckling. The critical buckling load is calculated from:

2

2crEIP

( k )π

= Equation 7.7

where E is Young’s modulus of elasticity, I is the moment of inertia of the cross-sectional area, is the length of the rod, and k is a constant that depends on clamping method (k = 1 for

pinned-pinned clamping method). The diameter of the rod that was selected was 0.06 m. The final mass for these selected dimensions for the I-beams and the rods was about 300 kg. See Appendix A.7 for detailed analysis. 7.7 – Module Docking Design Module docking is an integral part of the system design. Docking fulfills the system requirements of modularity, upgradeability and building a reconfigurable system. The Odyssey system will use a semi-autonomous docking process; differentiating from the usual practice of a mostly ground station or astronaut controlled procedure. The design requires the development of several unique connection mechanisms and the inclusion of essential interfaces. The design and approach to docking should accomplish the task and fulfill system requirements.

7.7.1 – Scope, Requirements, Choices, Design Methodology and Approach The system requirements state that the components must be upgradeable and modular; this creates the need for docking of segments and modules. The removable components are housed in modules connected to the segment core. In satisfying the upgradeable requirement the core is capable of accepting new modules and releasing old modules. The new modules are brought to the system through a docking process.


Final Report 78

The docking process produced obstacles concerning connection interfaces, docking mechanisms, module delivery methods, and a logical docking sequence must be developed. The modules contain no self supporting systems of power and housekeeping functions; therefore the connection must provide power and data interfaces on the module to code docking mechanism. The core to core docking mechanism must provide a fuel interface as well as the data and power needed on the module to core mechanism. The docking mechanism must provide a structural and hard-dock connection, resulting in the inclusion of a secure locking feature. The method of delivery must transfer the module from its orbit insertion point to the system and provide the propulsion necessary for docking. Finally, the docking process must follow a logical sequence of steps to achieve docking. In the docking mechanism design the first design choice was the selection of the interfaces present on the connection faces. These connections provide the power connection, data/information transmission and fuel transfer when necessary. The first choice made involved the selection of internal data transfer. A wireless network was selected to achieve the data transmission needs to reduce connector complexity and system mass. The unique needs of power and fuel transfer dictated their respective designs. Power is transferred in a simple connection consisting of two tabs to create a circuit; this is expanded for larger connections. The fuel interface consists of a two line connection for redundancy. The propulsion system included values to control the flow across the connection, with seals incorporated on the interface to prevent leakage across the connection when the line is pressurized.

7.7.2 – Concept Selection The development of the docking concept required investigation into several module delivery methods, docking sequence and connector and mechanism concept design. The mechanism development evolved through the need to provide necessary interfaces through the docking connection. The docking mechanism must also provide the hard docking and structural connection. The need for a structural connection arises from the presence of modules attached to the core during launch and the desire to prevent the connection from failing due to the forces experienced. Another purpose of the mechanism is to provide the interfaces needed for power and data transfer systems. The first concepts consisted of a single probe captured by means of a robotic arm or deployable tethers to “reel” in the module or segment. Both methods provide little orientation control of the docking object. The lack of control made it difficult to align connectors on the interfaces, and was the basis for elimination. The deployable tethers were also eliminated because they might become entangled when deployed, or miss the intended target resulting in the need for several attempts, lengthening the docking process. Initially a hard data connection was considered; this was eliminated to simplify the connection and reduce weight by incorporating a wireless system only requiring a small hole through which the signal can pass. The docking process evolved from the consideration of various delivery concepts. One initial concept consisted of using a small propulsion unit to bring the docking object into the range of a capture device. This concept generated design problems in the development of the capture device. A robotic arm of 3 m or more adds a significant amount of weight and is complex in operation.


Final Report 79

Another option considered was the use of a retriever nanosatellite flying in formation with the main satellite system. The retriever nanosatellite was eliminated because it increases the amount of fuel the system uses, by 25% if used for retrieval, de-orbit, and to perform orbit keeping maneuvers. The retriever also required designing a refueling port capable of mating with the satellite propulsion lines and removing fuel. An option for de-orbit was also considered. The de-orbit scheme considered used a small solid rocket motor to de-orbit the used module when it was replaced. In order to de-orbit a spacecraft module from our 800 km circular orbit, 220 m/sec of Δv must be supplied. Typical internal specific impulses of solid rocket motors are 280 sec. For a module mass of 500 kg, the maximum expected, at least 40 kg of solid propellant are required. Solid rocket motors such as the Star 17, currently in production, can meet this requirement. Motors such as these cut the usable volume of the module down considerably, if they fit at all. The Star 17 is 17 in in diameter. The selected delivery method is the product of collaboration with the propulsion engineers. The delivery method consists of a Module Propulsion System (MPS) to deliver the module to the satellite and is attached through the entire process. The MPS is a passive system responding to only the commands of the satellite and ground station through the docking process. It is a disposable unit consisting of clusters of thrusters and able to provide low levels of thrust for the final docking stages. The MPS is able to de-orbit an old module by attaching to the top of the module and carrying to a naturally degrading orbit. Incorporated into the design is a spring type mechanism to push the module radially away from the core.

7.7.3 – Discussion of Design Choices Design choices made at several levels led to the development of the final concept by influence of system requirements or by the concept progression. A major design selection concerned the mode of data transfer with the main computing system and the components in each module. The two options considered were a wired system or a wireless system. A wired system is traditionally used in satellite data transfer and processing, but for a satellite of Odyssey’s magnitude and expandability the projected mass of wiring approached 6% of the system dry mass. The desired wireless system is in the IEEE 802.2 family, but has not yet been fully developed or tested in a space-like environment; considering the current rate of wireless technology development it was assumed the technology would be mature at the time of satellite construction. The development of docking sequences provided several obstacles to overcome. The first design choice made was the use of semi-autonomous docking vs. fully autonomous vs. ground controlled docking. Fully autonomous docking relies on the perfect performance of preprogrammed instructions. The instructions are limited to the task and cannot handle unexpected events because they cannot solve problems when they arise or perform accurate progress checks during the docking process; making it impractical for this application. In ground controlled docking, there are no autonomous processes because the ground station dictates all actions. The benefit of ground control is the constant monitoring that prevents docking related problems such as misalignment and the option of correcting problems. Docking guided by ground control alone is a longer process because of the delay in information transmission


Final Report 80

between the satellite and ground station, this delay and the need for a full mission control team in the ground station make it an unlikely option. The selected option of semi-autonomous docking provides the benefits of previously discussed options. Semi-autonomous docking concept separates the process into a series of actions followed by a checkpoint. The actions consist of a short set of instructions performed autonomously by the satellite, and then a signal is sent to the ground station for approval and the next set of directions. The disadvantage to semi-autonomous docking is that it relies on the ground station for approval at key steps, lengthening the process. It is more successful than autonomous docking and shorter than ground controlled docking.

7.7.4 – Module Docking One key function of the satellite is the docking of the module to the segment core. It is through docking that the satellite becomes easily upgradeable and able to recover from damage to components. The docking concept developed is based on the concept presented in the paper Concepts and Technology Development for the Autonomous Assembly and Reconfiguration of Modular Space Systems by Lennon Rodgers and David W. Miller.

7.7.4.1 – Docking The docking process consists of a collaborative effort between the structure and propulsion subsystems. The obstacles to overcome in the docking connection are providing the required interfaces and to provide a structural connection. The structural connection provided by the docking port transfers loads and supports the module or core in several ways. The modules on the optical segment are present at the time of launch; therefore, they must support the module through the axial load of 7g. The docking mechanism also provides the hard-dock connection, a physical attachment between two components of a system. The final concern in docking is solving the problem of ensuring the proper alignment of the docking mechanisms. The docking mechanism is designed to withstand launch forces by transferring the load from the modules to pins connected to the core. The core connection points then transfer these forces to the load bearing structure, and ultimately to the launch vehicle. The module connector has two pins through which the load is transferred. These pins were sized according to the expected maximum mass supported through launch to prevent yielding of the material. The resulting radius of the circular cross section pin was doubled to account for cases when the module may contain more mass or if one pin is defective. The connector pins provide the hard-dock interface when coupled with a motorized locking plate. The locking plate serves two purposes; it protects the wireless port from the contamination and the space environment and it slides down behind the head of the connector pin. When the plate slides to the fully engaged position it securely locks the pin in place. Each connector has one pin and one locking mechanism, allowing for the design of a universal connector, only requiring a rotation angle of 180o for the mating connector. The connector is shown below: the outer face shows the pin extending from the surface and the reverse shows the docking motor. The locking plate mechanism is also shown in Figure 7.11.


Final Report 81

Figure 7.11 – Alignment of Docking Mechanisms on Final Docking Approach

When the module is within 3 m during the docking process a laser and camera perform measurements and alignment on the fine-tuning module orientation. This monitoring is done to ensure the proper alignment of the docking connectors. The approaching module is passive during this approach and receives necessary directions from the system to complete the process. The docking process is explained in the attitude and control portion of the paper, section 8.7. The docking mechanisms also include a spring-like mechanism to aid in undocking. The process is carried out by a paraffin actuated mechanism to provide the initial physical force to remove the module from the bay. The wax is heated for the compression during docking and reheated to eject the module at the end of its life. This mechanism is reusable and will be incorporated on the core external surface.

7.7.4.2 – Interfaces The need to incorporate connections to power, data transfer, and fuel systems produced a design problem. The use of several plug-in type interfaces on the docking connection results in the need for a docking process to be accurate to less than 1 cm. This is difficult to achieve, even under ideal conditions, because most measurement devices cannot accurately track to this degree of precision. Therefore, the interfaces needed must be simple and forgiving of slight misalignments during the docking procedure. The power interface is achieved through the incorporation of two metal tabs on the external surface of the connector. The metal tabs are exposed to the external environment during the docking process and come into contact at the last docking stage. The use of two tabs was desired to create a closed circuit. Using metal tabs to transfer electrical power simplified the design because it allows for slight misalignments to occur and is still operational. The tab design also eliminated the need for plug type interfaces on the surface of the connector reducing the risk of damage to the connector through breaking during docked launch or space transit for a single module.

Locking Plate

Connector Pin


Final Report 82

The data connection is achieved through the incorporation of a wireless network. The wireless network reduces the amount of internal wiring, and simplifies the connector by reducing the number of physical interfaces. Since the connectors are metal and the satellite structure is metal, the wireless signal must be routed through an open port or hole. This port is provided by a cutout through the docking mechanism plates, allowing the wireless signal to pass through the connection. The port is covered prior to docking operations by the locking plate. The locking plate has an extension over the hole to prevent damage from contamination during launch and in space. During the last stage of docking when the locking plate slides into place it uncovers the wireless signal port. Once sufficient power is supplied when the tabs are in contact the data transmission and reception begins. See Figure 7.12 for a depiction of the interfaces.

Figure 7.12 – Docking Mechanism, Front View

7.7.5 – Segment Docking Concept The docking concept for segment connection is based on the module docking process and mechanisms with design alterations to accommodate the changed needs. The difference in the segment docking mechanism or the core to core connection, is the inclusion of fuel line connectors and an increased number of power connections. Also including more connector pins to ensure a secure connector on a larger structure. The segment docking is achieved through a core to core connection. The connection provides the various interfaces needed, one priority is the structural connection. The core connection required a more structurally secure connection. The propulsion segment and the optical payload segment will be attached at launch; therefore this attachment was designed to handle this loading. The connection is also larger on the core connection vs. the module to segment connection, so the hard-dock connection became more robust. The hard-docking, or physical dock, was accomplished through the use of four connector pins. The use of four connector pins and locking mechanisms ensured that the core connection disperses the forces. The power interface, like the structural interface is doubled. The doubling of the power tabs allows for more power transmitted across the connection. The core requires more power since it must relay the power to all of the modules and the wireless network nodes located in the core.

Connector Pin

Wireless Signal Port

Power Tab

Connector Pin Interface


Final Report 83

The wireless transmission port remains the same size from core to core. The wireless network will have a node at that point collecting and relaying the data to and from the main connector across the interface. Core docking also requires the inclusion of a fuel interface for the extension of the fuel lines. The extension is necessary to transport the fluids of the propulsion subsystem to the thrusters at the end of the optical segment. This connection is also necessary for possible core expansion in the future. Further expansion of the core beyond a few meters will require additional thrusters to maintain proper orientation. There are two hydrazine line interfaces, for the sake of redundancy in the case of a leak in one line. This connection is controlled by valves designed by the propulsion subsystem to regulate the fuel flow, and by a series of seals and fasteners to prevent external leaks across the interface. The same solution is applied to the hydrogen connection necessary for fuel line pressurization. Two figures follow depicting the connection mechanism with labeled interfaces.

Figure 7.13 – Core Docking Mechanisms, Front View

The segment docking procedure is similar to the module docking procedure. It follows the same steps of bringing the new part to system, in this case attaching a new core segment. The segment is propelled with a smaller version of the main propulsion unit, which is detachable and utilizes the vacant connector at the end of the segment. This unit may stay in place until it is necessary to discard it. The satellite will locate the new segment using antennae and LADAR tracking. It also uses laser fine positioning methods and cameras to ensure proper alignment. When properly aligned the new segment is pushed into place and the core connector locks together, creating the hard-dock connection. The interfaces complete their connections via tabs (power), sealing mechanisms and valves (fuel lines) and activating the wireless network nodes. This process is also subject to the ground station approval at critical stages through checkpoints mentioned in the module and segment docking procedure. 7.8– Cost Estimation As explained in Section 10.1 about cost analysis, the detailed list of parts and their costs is beyond the scope of the cost analysis in this project. Costs are estimated based on cost-per-mass

Power Tabs

Hydrogen Interface

Hydrazine Interface

Connector Pin Connector Pin

Interface


Final Report 84

basis from historical data found in Space Mission Analysis and Design.[7.9] The final cost of the structures and thermal subsystems was estimated to be $5,150,000. See Section 10 for more details. 7.9 – Summary and Lessons Learned The structures and mechanisms subsystem is responsible for providing a structurally sound structure for launch survival and docking operations throughout its life time. This was achieved through the development of the core and module structures, and through the development of the propulsion segment structure. Additionally, the conceptual docking mechanisms were developed with the process development in collaboration with the propulsion subsystem. The development of the respective structures has a limited about of fidelity because lab testing of structural elements is not possible in the class structure and time frame. The docking scheme is of a low fidelity level because it is a conceptual design and does require testing before it is integrated into a system. Through the structure and mechanism development it shows that modularity increases the structural mass of a system. This is impractical for space systems unless it is essential to the mission for something such as docking. Docking also complicates the design by requiring the development of specialized mechanisms and processes.


Final Report 85

7.10 – References [7.1] Sarafin, T. P., Spacecraft Structures and Mechanisms, Space Technology Library, 1995.

[7.2] Conley, P. L., Space Vehicle Mechanisms, John Wiley and Sons, 1998. [7.3] Cook, Robert D., David S. Malkus, Michael E. Plesha, and Robert J. Witt. Concepts and

Applications of Finite Element Analysis. 4th ed. Madison: John Wiley & Sons, 2002. 19. [7.4] Matweb Material Property Data. 4 Nov. 2006 <http://www.matweb.com>. [7.5] "Pyrolytic Graphite." GE Advanced Ceramics. 4 Nov. 2006

<http://www.advceramics.com>.

[7.6] Peery, David. Aircraft Structures. New York: McGraw-Hill, 1950. 279. [7.7] Sun, C. T., Mechanics of Aircraft Structures, John Wiley and Sons, 1998. [7.8] Hibbeler, R. C., Statics and Mechanics of Materials, Prentice-Hall, Inc., 1993.

[7.9] Larson, Wiley J., and James R. Wertz, eds. Space Mission Analysis and Design. 7th ed.

El Segundo: Microcosm, Inc., 2005.


Final Report 86

8.0 - Attitude Control System The attitude control system maintains the spacecraft’s orientation in orbit. A combination of momentum wheels and maneuvering thrusters apply torques to the spacecraft to maintain or change the orientation. It is vital to have an attitude control system that can control the spacecraft to the accuracy required by the scientific instruments. 8.1 - Scope and Purpose The primary purpose of the Attitude Control System (ACS) is to determine and control the vehicle attitude. This section discusses the design of the ACS onboard the Odyssey. It covers the evolution of the design from its inception to the final concept. First, the functional requirements are presented. Next, the design challenges as well as the various ACS types are considered. The required hardware choices are then discussed. The ACS mass, power, sizing, and location requirements are given. Finally, the ΔV budget is presented and discussed. 8.2 - Tasks, Functions, Requirements and Design Methodology There were three main tasks in the ACS design process. These included determining the appropriate ACS type to use aboard the Odyssey; determining the hardware to be used and its specifications; and determining the ΔV budget. Another important task was to coordinate with the propulsion group. The propulsion and ACS functions and hardware closely relate to one another, and in some instances overlap. The ACS thrusters, for example, will be mentioned briefly in this section, but will be discussed in further detail in the propulsion section. There are four main functional requirements for the Odyssey’s ACS. These include determining and controlling the vehicle attitude, implementing changes in orientation when required, maintaining pointing accuracy of extensions, and aligning modules and segments for on-orbit rendezvous with the satellite. Each of these functions must be performed by the ACS. According to the system requirements document, the primary mission of the Odyssey is observation. As a result, the system is required to obtain Hubble optical capability with a resolution of 0.05 arcsec. This was the primary driver for the design of the Odyssey’s ACS. This is discussed in section 8.3.3. Furthermore, the system requirements state that gyroscopes are to be used to determine vehicle attitude, reaction wheels are to be used to control vehicle attitude, and ACS thrusters are to be used to unload reaction wheels. These design choices are discussed further in section 8.4, as well as the selection of the Fine Guidance Sensors (FGS). The design methodology consisted of six steps:

7) Determining system requirements 8) Identifying the functional requirements of the ACS 9) Choosing type of ACS to be used 10) Studying systems with similar pointing requirements 11) Identifying required hardware


Final Report 87

12) Determining the ΔV budget 13) Determining mass and power requirements for the ACS hardware

8.3 – Choices - Decision Justification and Design Drivers The three design drivers for the attitude control system were disturbance torques, three-axis stabilization, and center of mass of the system. These each needed to be fully investigated to define the system.

8.3.1 – Disturbance Torques In orbit, the ACS must compensate for different environmental torques. These include:

1) Drag torque 2) Gravity gradient torque 3) Magnetic torque 4) Solar torque 5) Spacecraft generated torques

Drag torque is the dominant cause of orbit decay. At an altitude of 785 km, this torque is small, but it varies greatly with atmospheric density and solar activity and cannot be ignored. The ACS thrusters must perform periodic orbital maintenance maneuvers to main the required altitude. Gravity gradient torque occurs because the lower portion of the satellite is subjected to higher gravity forces than the upper portion. Magnetic torque is due to the magnetic field of the Earth; it decreases as the altitude of the orbit increases. Conversely, the solar torque increases in relative importance with increasing orbital altitude. [8.1] Torques are also generated aboard the satellite. For example, a torque is generated if the thrusters fire along a line that does not contain the center of mass of the system. The ACS must be designed to compensate for these different torques. The Odyssey’s center of mass is presented next. The specific ΔV requirements are discussed in section 8.7.

8.3.2 – Center of Mass Figure 8.1 defines the Odyssey’s body axes. The primary axis, V1, runs through the core, V2 parallels the solar array blankets, and V3 parallels the space viewing telescope (SVT). The center of mass of the Odyssey was found to be CM = -2.51 V1 + 0.03 V3 [m]. This was measured from the end of the optical segment that connects to the propulsion segment. While there is a slight deviation of the CM off of V1, it is still located inside of the core. The torque generated by this deviation will be small, and will not contribute significantly to the ∆V budget. Center of mass calculations can be found in Appendix A.8.


Final Report 88

Figure 8.1: Satellite Body Axes

8.3.3 – Three-axis Stabilized System A three-axis stabilized system is defined as one which “actively maintains the vehicle axis system aligned with a reference system.”[8.1] The selection of a three-axis stabilized system was a direct result of the pointing requirements for the SVT. Pointing accuracy requirements vary for different components aboard satellites. Typical requirements for solar arrays are 4° to 10°, for high gain antenna are 0.1° to 0.5°, and for optics, telescopes and cameras are 0.001° to 0.1°.[8.1] From this, it is clear that the design driver for the ACS is the optical instrumentation. Table 8.1 shows the pointing accuracy of common ACS types.

Table 8.1: Pointing Accuracy of Common ACS Types[8.1]

Gravity Gradient Spin Dual Spin Three Axis

Momentum Bias

Pointing Accuracy 5° 1° 0.1° 0.001° 0.1° to 3° A three-axis stabilized system is required to achieve the pointing accuracy requirements of the SVT. This choice provides significant benefits. The pointing accuracy is limited only by the precision of the Odyssey’s sensors. Of all of the ACS types, it is the most adaptable to changing requirements. This choice allows Odyssey to adapt to new system requirements. However, this selection comes at a cost. The required hardware is complex, heavy, and expensive. 8.4 – ACS Hardware HST’s pointing control system (PCS) was used as a baseline for the Odyssey’s ACS because of the similarities between the two systems. However, the space shuttle provides orbital


Final Report 89

maintenance to HST, while the Odyssey will have to provide its own. From this and other ∆V requirements, it was determined the Odyssey would also require a set of ACS thrusters. These are discussed in greater detail in the Propulsion section. All of the ACS hardware, with the exception of the ACS thrusters, will be located on the optical segment. Table 8.2 traces the hardware to specific function requirements of the ACS.

Table 8.2: Traceability of ACS Hardware to Functional Requirements Functional Requirement ACS hardware (1) Determine vehicle attitude Rate-sensing gyroscopes (2) Control vehicle attitude Reaction wheels, ACS thrusters (3) Implement desired changes in orientation Reaction wheels, ACS thrusters (4) Maintain desired pointing accuracy of SVT Fine Guidance Sensors (FGS) (5) Align modules and segments for on-orbit rendezvous with satellite

ACS thrusters

8.4.1 – Rate-sensing Gyroscopes A rate-sensing gyroscope measures attitude rate motion about its axis, and provides the attitude reference when maneuvering the satellite.[8.2] A minimum of three gyroscopes are required for determining the vehicle attitude; one for each of the primary axes. An additional three gyroscopes were included for redundancy for a total of six. Like on the HST, the gyroscopes are housed in pairs in Rate Sensing Units (RSU). They are located aboard the optical segment in Modules 1.1, 2.1, and 3.1. They weigh approximately 5.5 kg apiece, including housing. Figure 8.2 is a picture of the rate-sensing gyroscopes used aboard HST. [8.2]

Figure 8.2: HST’s Rate Sensing Gyroscopes[8.3]

The gyroscopes require periodic updates to remove random drift. The gyroscopes have a drift rate of 1.4 ± 0.7 mas/sec.[8.3] The FGS, discussed in Section 8.4.3, provide star reference data to realign the gyroscope, to maintain the accuracy of the measurements over time.


Final Report 90

8.4.2 - Reaction Wheels Reaction wheels move the satellite into required orientations, and compensate for disturbance torques. As with the gyroscopes, one reaction wheel is required for each of the primary axes. A fourth reaction wheel, in a position oblique to all axes is included for redundancy. This is a common practice. In this way, any single reaction wheel failure can be accommodated by the redundant wheel.[8.1] The reaction wheels are housed in pairs in Reaction Wheel Assemblies (RWA).[8.2] They are located aboard the optical segment in Modules 1.5 and 3.5. Figure 8.3 shows a picture of the HST’s RWA.

Figure 8.3: HST’s Reaction Wheel Assembly[8.2]

The reaction wheels were selected to be HR195 manufactured by Honeywell.[8.1] The reaction wheel characteristics are given in Table 8.3. The reaction wheels can reorient the satellite 90 ° in 14 minutes.[8.2] This selection was based on the mass of the Odyssey and its maneuvering requirements.

Table 8.3: Reaction Wheel Characteristics[8.1] Reaction wheel HR195 Angular momentum 265 N-m-s Output torque 0.7 N-m Peak power 400 W Power holding max speed 45 W Wheel speed 3000 rpm Weight 48.0 kg Operational temperature -18 to +49 ºC Motor type DC

8.4.3 - Fine Guidance Sensors In the early stages of the design process, star trackers were included in the attitude control system to correct errors in the gyroscope readings. After further investigation into HST, it was determined that star trackers would be insufficient to obtain the required pointing accuracy of the SVT. Because a pointing accuracy of 0.05 arcsec is required, this portion of the ACS had to be redesigned. Instead of star trackers, three Fine Guidance Sensors (FGS) were incorporated into the design. FGS were first developed for HST.[8.5] They provide data to the spacecraft’s ACS,


Final Report 91

and they act as a science instrument. The three FGS are located on the super module at 90º intervals around the base of the SVT.[8.2] They weigh approximately 220 kg each. Figure 8.4 shows a cutaway view of FGS.

Figure 8.4: Cutaway View of FGS[8.2]

8.5 – Mass and Power Requirements and Discussion of Cost The mass and power requirements for the attitude control system hardware are given in Table 8.4. The sizing of the hardware is given in Table 8.5. In total the ACS weighs approximately 980 kg, and consumes approximately 430 W of power.

Table 8.4: ACS Mass and Power Requirements Mass, each Power, each Number Mass, total Power, total RSU 11 kg 40 W 3 33 kg 120 W Reaction wheels 48 kg 60 W 4 192 kg 240 W FGS 220 kg 20 W 3 660 kg 60 W Thrusters 8 kg 1 W 12 96 kg 12 W TOTALS: 981 kg 432 W


Final Report 92

Table 8.5: ACS Hardware Sizing

Height (m) Width (m) Length (m) Outside Diameter

(m) RSU (contains two gyroscopes)[8.2] 0.3 0.3 0.2 --- Reaction wheel[8.1] 0.5 --- --- 0.7 FGS[8.2] 0.5 1.0 1.5 ---

Accurate cost estimates for the rate-sensing gyroscopes and HR195 reaction wheels were not available. The three FGS are estimated to cost $63 million.[8.5] The FGS are heavy, complex, and a relatively new technology. This will be the bulk of the ACS cost. 8.6 – Docking Process To fulfill the requirements of modularity and upgradeablity, docking became a necessary operation. The module docking process is the process in which most of these upgrades will occur. The process will consist of docking a new module with update components and undocking the old module. A docking concept was developed for the modules; the process can be expanded to accommodate the docking of segments.

8.6.1 – Locate Module After the module is launched into orbit with the system, the first step in the docking process is to locate the module. The module is located by the frequencies transmitted by the antennae of the Module propulsion system (MPS). Each of the antennae emits a unique frequency, which is also used in the initial orientation and range measurements. Once located the system directs the MPS movements through the docking process. The arrangement of the antennae on the MPS allow for the module to attain a desired approach orientation. The orientation is attained by the spacecraft and ground team searching for the desired relative distances between the sources of the antennae frequency. The thrusters on the MPS are then fired to achieve the correct orientation. The module is also tracked by the use of Laser Detection and Ranging or LADAR once it is in a 3 km radius of the system. The incorporation of LADAR provides many benefits to the module tracking. LADAR allows for the tracking of the range, velocity and orientation of the object, in this case an approaching module. The orientation is available through a visual image, or LADAR signature. The geometry of the module and the MPS will provide several unique signatures based on orientation. The images from each scan are analyzed to check through a library of stored orientations. It will identify when the correct orientation is reached and perform checks on each subsequent velocity and range scan to ensure that the orientation is maintained. The ability to produce an image reflecting the contours of a 3-D object greatly aids in simplifying the task of fine orientation later in the docking process.

8.6.2 – Direct Module The system receives the signals from the antennae and the LADAR readings. These are processed by the main computing system to determine the position and orientation of the module


Final Report 93

relative to the satellite. The data is then fed into ranging and control algorithms, incorporating Kalman filtering to determine the set of instructions to be sent to the MPS dictating the next series of movements. The process of tracking and instruction relay continues through the docking process.

8.6.3 – Dock Module The module continues its approach to the system by the location and direction methods described until it reaches the final approach stage. At the final approach the module is within 3 m when the laser fine positioning and the visual cameras track the progress on the plane of the docking mechanism. The system collects the visual camera images and compares them to reference images of the correct orientation. When the correct orientation is reached it is continuously monitored and maintained by the system. The lasers lock on to specific points on the bottom surface of the module and track the point location to ensure an alignment accuracy of within centimeters. The docking mechanism alignment pin interface is angled to allow for misalignment on the degree of a few centimeters. The MPS continues to push the module in until the connector pins are locked in place completing the hard-dock of the system. At the completion of the hard-dock process the power tabs are in full contact; therefore powering the modules. Once powered, the module is able to transmit data signals via the wireless network through the wireless signal port. The signal port is opened when the pins are locked in place by the locking plate sliding into the locked position. When the module and segment are hard-docked with the interfaces complete the docking process concludes. At the conclusion of the docking process the MPS detaches and is free to de-orbit itself, or to remove and de-orbit an out of service, or old module. 8.7 - ΔV Budget The ΔV budget is given in Table 8.6. The orbit insertion accuracy of the launch vehicle is ±10 km altitude and ±0.1° inclination. Using this information, the ΔVmax required to correct the orbit insertion error is 18.7 m/s. Furthermore, periodic orbit maintenance maneuvers will also require small ΔV’s from the thrusters. It is necessary to carry enough fuel aboard to perform a deorbit if necessary. As discussed in section 8.3.1, the ACS will also have to compensate for different environmental torques. At this point in the design process, there is not enough system definition to determine these values accurately. Even with complete system definition, these values vary greatly. For these reasons, the budgeted ΔV for the environmental torques is an estimate. Detailed ΔV calculations can be found in Appendix A.8.


Final Report 94

Table 8.6: ΔV Budget ΔV Reference Orbit insertion* 0.52 (m/s)/km Calculation Orbit maintenance 1 (m/s)/yr Calculation Deorbit [8.6] 208.3 m/s SMAD Environmental torque compensation

5 (m/s)/yr Estimate

Plane change** 135 (m/s)/º Calculation TOTAL * ΔV not provided by launch vehicle **Plane change ΔV listed for reference 8.8 - Summary – Lessons Learned and Future Efforts The design of the ACS is an involved process. It may be advisable for future design projects of this nature to assign more than one person to the ACS design. This will allow greater system definition. The ACS design to this point has selected and sized the major hardware elements. The ΔV budget has also been determined. Future work must include defining the control law and attitude determination method.


Final Report 95

8.9 - References [8.1] Dukes, E.M. “Part 5: Attitude Control.” Wren Software Inc. 4 August 2001. [8.2] “Hubble Space Telescope: Servicing Mission 3A Media Reference Guide”. 27 December

1999. 22 October 2006. Prepared by Lockheed Martin. <hubble.nasa.gov/a_pdf/news/SM3A-MediaGuide.pdf>.

[8.3] “Servicing Mission 3A Overview.” 27 December 1999. 22 October 2006.

<http://sm3a.gsfc.nasa.gov>. [8.4] “Hubble Space Telescope Primer for Cycle 12”. Space Telescope Science Institute. 1

November 2006. <http://www.stsci.edu/hst/proposing/documents/cp_cy12/ Ch_ 3_TelescopePer3.html>.

[8.5] “FGS Instrument Handbook.” Space Telescope Science Institute. 15 November 2006.

<http://www.stsci.edu/hst/fgs/documents/instrumenthandbook/>. [8.6] Larson, Wiley J., and James R. Wertz, eds. Space Mission Analysis and Design. 3rd ed. El

Segundo: Microcosm P, 2005.


Final Report 96

9.0 – Launch Vehicle System To provide the ΔV necessary to achieve orbit, a launch vehicle system was chosen. This choice was based on the mass, volume and cost of a gamut of available launch vehicles. Current launch vehicle systems are very costly and an attempt was mad to choose a low cost, reliable system. 9.1 Scope and Purpose This section entails the selection process, technical specification, and integration of the Odyssey’s launch vehicle system (LVS), and the site from where it will be launched. Each part of this section details the role played by the LVS and its specific relation to the Odyssey spacecraft. These sections contain specific information regarding the functions, trade studies, design drivers, decisions, launches, launch site, and launch timeline resulting in the chosen LVS. The purpose of the Odyssey LVS is to provide the spacecraft with all the necessary components to place the spacecraft in the desired orbit. The LVS will deploy the spacecraft within an orbit accuracy range dependant on the specific launch vehicle. The ability of the launch vehicle to place the spacecraft in the designed orbit is based on the altitude and inclination of the orbit, spacecraft’s weight, and the launch vehicle’s capabilities. Not only does the LVS provide transportation to the desired orbit, but it also protects the spacecraft from the environment until it is deployed. While the spacecraft is being launched into its orbit it will be exposed to acceleration forces, temperature ranges, and acoustics. While the Odyssey is designed to meet the launch vehicle requirements the payload fairing provides protection from the atmosphere. Also included in the LVS are the ground operations required to prepare the system for its insertion into space. The Odyssey launch site is an important element in the progression of the spacecraft’s mission. The handling and protection of vital spacecraft systems before they are activated in the orbit environment are provided by the launch facility. Areas such as cleanliness levels, environmental controls, and launch vehicle integration must be met by the launch site. In case the initial LVS cannot meet the needs of the Odyssey, and in order to provide commercial competition, a backup launch vehicle was chosen that currently meets all Odyssey mission requirements. The backup launch vehicle becomes necessary in any schedule conflicts with the primary LVS or any other unforeseeable problems. 9.2 Tasks, Functions, Requirements, and Design Methodology The LVS of the Odyssey project has one basic requirement that all other tasks and functions fall under, and it is that the system be launched by 2011. There are several other stipulations on how the system should be launched, but the basic driving requirement is the time when it occurs. Section 2 of this document outlines the system requirements for the LVS. There is one main tier one requirement, six tier two requirements, and six more tier three requirements that the LVS must meet.


Final Report 97

To meet the tier one requirement of having a launch by 2011, all current launch vehicles that were planned to be phased out in the next five years were eliminated. Then the launch history of each vehicle was reviewed and all vehicles not having multiple launches in a year were eliminated. Requirement 6.1 is satisfied because no LVS that has not been tested or will have had a successful launch by 2011 was considered in the selection process, giving it a readiness level of seven or higher. Although requirement 6.2 sounds exactly like the tier one requirement, it implies that there will be multiple launches during the mission history, and the LVS must meet this initial deadline. This deadline is driven by two requirements 6.2.1 and 6.2.2, which state the LVS have the launch site and vehicle availability, along with the production capability to have another launch in the mission time frame. These requirements were met by determining if there were any launch restrictions on the launch site, and what the expected launch and production cycles were for the LVS. There are no restrictions on the launch site in terms of schedule delays and the LVS production cycle is met by all considered launch vehicles both primary and backup. The four requirements under section 5.3 focus primarily on the specific insertion accuracies of the LVS being used. In the Odyssey mission the primary launch vehicle will be multiple versions of the SpaceX Falcon 9 launch vehicle. The Falcon 9 provides insertion accuracies equal to and possibly greater than the required values. The Falcon 9 uses redundant guidance systems and an additional GPS based system to give it these accuracies. These requirements were driven by the limits of the propulsion systems used to deliver the different modules to the existing spacecraft. The backup LVS, Atlas V, also meets the insertion requirements. The LVS must provide a ΔV of greater than 9.2 km/s in order to meet the orbital requirements, as stated in 8.1.1. Since both the primary and backup launch vehicles are based on an Evolved Expendable Launch Vehicles (EELV) design the LVS can provide a range of required Δv for a large range in mass of the spacecraft. The only action required is to meet the requirement is to match the spacecraft’s mass to the specific version of the LVS. The requirements of section 5.3.3.1 are focused on the limits of the launch site used and its inclination capabilities. Due to locations of populations and landmasses and the dangers of a launch explosion, not all launch sites can be used to launch into the desired 98.5° inclination. All launch sites not able to meet the desired inclination were removed from the possible launch site list, and all launch vehicles that could only be launched from the removed sites were eliminated as well. Finally, the last requirement was imposed to set limits on the size of the spacecraft and what LVS were needed to meet the sizing needs. Since the largest fairing size on common EELV designs were 4.6 meters it became the maximum spacecraft diameter and the required fairing size for the primary and backup system. With the Falcon series and the Atlas V both being EELV based designs, they had a standard fairing meeting this requirement.


Final Report 98

The selection method used in selecting the LVS is a combination of the S.M.A.D. [9.5] outlined process, and an informal process generated as new system information and analytical fidelity increased. Table 9.1 represents the information used in the S.M.A.D. process. Because the Odyssey will never be built and tested, and because the analytical fidelity of the system is limited, certain aspects of the selection process were modified to fit this specific project. Areas such as acoustic frequencies could be found for the launch vehicle, but in the current design it would be difficult to test the system to see if it met the criteria. The overall design method was driven by time based decisions and events occurring in the design process. After all current and possible launch vehicles were found no decision could be made until an orbit and system mass were provided. Once the orbit was chosen all LVS and launch sites not meeting the criteria were removed. Since three of the four systems left were EELV based designs nearly any system mass could be launched. This changed the focus of the selection process to a cost and reliability based selection. The Falcon launch series became the selected choice because of its cost and predicted reliability. Although it has not had a successful launch, the significant reduction in cost is such that three launches of a Falcon would equal one launch of any other LVS considered. This was the main focus of the selection process and although the formal design method wasn’t followed exactly, if the selection process was done again using Table 9.1 the same vehicle would have been chosen.


Final Report 99

Table 9.1 – Steps in Selecting a Launch System [9.5]

Step Comments and Required Information 1. Collect requirements and constraints, which depend on the mission operations concept. Consider the deployment strategy.

Number of spacecraft per launch Spacecraft dry weight Spacecraft dimensions Mission orbit Mission timeline Funding constraints

2. Identify and analyze acceptable configurations for the launch system.

Include the following information for each potential configuration: Weight of spacecraft propellant Orbit insertion stage weight if required Weight of booster adaptor Performance margin available Boosted weight capability Reliability

3. Select launch system for spacecraft design. During conceptual design, identify several potential launch systems to make the launch more likely.

Criteria based on the following parameters: Boosted weight capability Cost Performance margin available Reliability Schedule vs. vehicle availability Launch availability

4. Determine spacecraft design envelope and environments dictated by the launch system selected.

Include the following information for each launch system, and include the worst-case environments for combined launch systems: Fairing size and shape Maximum accelerations Vibration frequencies and magnitudes Acoustic frequencies and magnitudes Temperature extremes Air cleanliness Orbital insertion accuracy Interfaces to launch site and vehicle

5. Iterate to meet constraints on performance, cost, risk, and schedule.

Document and maintain the criteria, decision process and data to support program changes

9.3 Choices - Trade Studies and Decision Justification The main trade studies conducted in this section of the design mainly focused on comparing cost and reliability to the capabilities of the LVS. Because current launch vehicles base their payload integration around standardized metrics, integration of a payload design to any current launch vehicle would be nearly the same. Areas such as payload adaptors, thermal environments, loading forces, and cleanliness factors fall in a range which is almost identical in all current


Final Report 100

launch vehicles. Based on these factors cost and reliability, based on capability become the most important factors. The first trade study conducted was cost versus capabilities. Launch vehicles chosen to be compared were those that would still be active by 2011, and those meeting the necessary predicted launch capabilities. The launch capabilities of these vehicles fall within a variation range of about 1000 kg. Considering many of the vehicles are EELV’s the comparison range can be reduced based on the model of the system, this results in a more accurate comparison. Figure 9.1 shows the cost of possible launch vehicles for the Odyssey. Cost for the Proton, Ariane 5, and Soyuz are not exact price quotes from the manufactures, but cost history and industry comparisons give reasonable results. The Falcon 9 is cheaper than all systems by almost 70%, making it the best choice based on cost.

Launch Vehcile Cost (Similar Capabilites)

020406080

100120140

Falcon

9

Delta I

V

Atlas V

Proton

Ariane

5Soy

uz

Cos

t ($M

)

Figure 9.1 LVS Cost Comparison [9.3]

The second trade study conducted was reliability versus cost. After the first trade study was conducted the selected launch vehicles were compared based on their mission success rate. Figure 9.2 gives the reliability comparison for the considered vehicles based on total mission success. Falcon 9 is currently has no success rate because it has not conducted any launches. Although the Falcon 9 has not had any launches their reliability is predicted to beat existing competitors. This prediction comes from a historical analysis of launch failures and using redundant and reliable components in the subsystems that typically cause failures, such as stage separation and engine failure.


Final Report 101

Reliability Comparison

70.0075.0080.0085.0090.0095.00

100.00

Falcon

9

Delta I

V

Atlas V

Proton

Ariane

5Soy

uz

Succ

ess

Rat

e

Figure 9.2 LVS Cost Comparison [9.4]

After the initial trade studies were conducted the orbit for the Odyssey mission was determined. When analyzing the orbit and searching launch sites it was found that the Proton and Soyuz launch sites were not capable of the desired inclination. This is due to their geographical location regarding cities in close proximity and certain land masses, thus removing them from the selection process. The ability of the Falcon series to meet our system’s required capabilities, provide increased reliability, and decrease overall cost made it the best LVS. Along with satisfying all the previous needs, Falcon will also be able to meet the future launch schedule. The Falcon’s diverse series of launch vehicles also allows for an increase in capability for our system. If larger or smaller, than typical, payload mass are desired to be launched a different series of vehicle can be chosen. 9.4 Design Evolution The design evolution of the launch vehicle selection process was primarily dependant the mass and dimension of the Odyssey component. In earlier designs the entire Odyssey was around 8000 kg and fit into a standard 4 meter fairing, but as the optical systems evolved the design became larger and heavier. The required fairing size increased to 4.6 m as well as the total mass to over 11,000 kg. As the final design was chosen the number of launches required increased to three, using two different vehicles and two fairing sizes. Although there is a Falcon launch vehicle capable of meeting the required mass, the dimensions of the space telescope became the design driver for the number and types of launches. 9.5 Launch Site The launch site Odyssey will use for the 98.5° near polar orbit is Vandenberg Air Force Base (VAFB). VAFB is located 34.67° North and 120.62° West on the coast of California. The Falcon is launched from Space Launch Complex 3 West (SLC-3W). Telemetry data is received and processed at two permanent sites, the VAFB Telemetry Receiving Site and Pillar Point. Each site can process code modulation rates up to 5 Mbps, which provides telemetry to the customer and range safety. Processing facilities consists of class 100,000 clean rooms, capable of lifting 5 tons, and handling hazardous chemicals such as Hydrazine. [9.6]


Final Report 102

9.6 Primary Launch Vehicles and Environment There will be two initially planned launches to complete the Odyssey so it may begin its primary mission. These launches will put all necessary components into orbit for the space and Earth viewing missions. A third launch will add the experiment segment with or without experiment modules already attached at a later date. The first Odyssey launch will consist of a Falcon 9S5 launching the propulsion segment, Earth viewing module, and the power and systems needed to operate the spacecraft. The total mass of the payload is estimated at 10,540 kg while the Falcon 9S5 is capable of taking 12,400 kg to our desired orbit. There is currently limited information on the Falcon 9S5; the concept uses the main core of a Falcon 9 and two Falcon 5 cores as boosters. Although the arrangement of the components on the 9S5 is different all payload interfaces and environments are the same for all Falcon 9 models. The fairing size of the 9S5 is 4.6 m in diameter and 12 m in height, the larger of the two Falcon 9 fairings. The second launch will consist of the super-module and the propulsion unit that will mate it with the rest of the Odyssey spacecraft. The launch will use the standard Falcon 9 with the large 4.6 m fairing. The mass of the payload will be 6,489 kg while the capability of the Falcon 9 with the large fairing is 6,550 kg. The third launch used to add the experiment segment will use the standard Falcon 9 with the smaller 3.2 m diameter fairing. The segment will consist of the core, segment propulsion system, and any experiments currently designed for the spacecraft. The current mass for the payload is 4,905 kg with vehicle capabilities of 7,000 kg but as mentioned the payload mass could increase for batteries, solar cells, or experiments added. The Falcon launch vehicles use an aluminum monocoque truncated cone and is the standard EELV 1.575 m diameter at the payload interface. [9.2] The Falcon uses a Lightband separation system which is a non pyrotechnic spring launching device making it lighter, cheaper and reducing the shock load by 95%. All thermal and cleanliness requirements set by the thermal team are met by the launch site, and a filtered conditioned three micron air purge system connected to the fairing through hoses. The greatest launch loads expected at any point in the launch are 6 and -2 g’s in the axial direction, and +/- 2 g’s in the lateral direction. The insertion accuracies for the LVS are +/- 10 km with an inclination accuracy of 0.1°. Considering the worst case scenario based on distance from deployment to the spacecraft a ΔV of 0.1819 km/s is required to reach the Odyssey. [9.5] 9.7 Backup Launch Vehicle In any launch vehicle selection process it is necessary to select a backup launch vehicle in case of scheduling issues, any other conflict that may arise, and to promote industrial competition with the primary LVS. [9.5] The backup vehicle chosen for the Odyssey project is the Atlas V, based on its 100% success rate and range of payload mass capability. The payload adaptors and environmental conditions are nearly identical to the Falcon 9, with the exception of the launch


Final Report 103

loads being reduced by 0.5 g’s. The launch site used by the Atlas V at VAFB is the exact site used by the Falcon 9 SLC-3W, thus producing no changes in the location assembly. The first launch would use an Atlas V 503, while the second and third launches would use the 500 series if necessary. [9.5] 9.8 Cost Estimate The estimated cost of the three Falcon 9 launches is $51, $35, and $27 million dollars. The prices are all inclusive of launch range, third party insurance, and standard payload integration cost, which results in a total of $113 million. A replacement of any of the Falcon launch vehicles by an Atlas 5 will increase the total cost by at least $130 million. The estimated cost for the baseline Atlas 5 is $130 million, but bigger model types have increased costs with the 503 model estimated at $200 million. If all Falcons were to be replaced by the Atlas V launch vehicles the estimated cost would be $460 million, not including many of the fees SpaceX includes. 9.9 Summary – Lessons Learned and Future Efforts In conclusion, the SpaceX Falcon 9 launch system will be the primary vehicles used in the three projected Odyssey launches. The Lockheed Martin Atlas V is the backup vehicle with both systems meeting the required integration and environmental standards. One of the biggest lessons learned in this design effort is that standardization of launch vehicles allows spacecraft designs to be less restricted. If satellite companies know they can create any satellite they want and are only constrained to one set of integration standards, their launch vehicle choices do not become restricted. This promotes competition in the launch vehicle industry and can help reduce costs and increase reliability in the future. In terms of the overall project one lesson that can apply to any design task is to identify design drivers and see how the design is being limited. At several points in a design it is required to make decisions making drastic changes in the system. It is important to see how your design drivers play a role in the decision made and whether it limits the system or allows the system to generate the best possible design. The future efforts of the launch vehicle section are to adapt the payload delivery system so that the Falcon 5, and even the Falcon 1, could be used to send up individual modules. This could provide the customer with more options and even provide an individualized Orbital Express type capability for the Odyssey spacecraft.


Final Report 104

9.10 References [9.1] Atlas 5 Payload Guide. Lockheed Martin. San Diego, California: Lockheed Martin, 1999.

21-25. [9.2] Capozzoli, Peter. "Falcon 9 Payload Environment." Email to Nicholas Sochinski. 20 Nov.

2006 [9.3] "International Expendable Launch Vehicle Data for Planetary Missions." NASA. 14 Sept.

2006 <http://www1.jsc.nasa.gov/bu2/ELV_INTL.html>. [9.4] "Space Launch Vehicle Reliability." 18 Oct. 2004. The Aerospace Corporation. 21 Sept.

2006 <http://www.spacex.com/>. [9.5] Space Mission Analysis and Design. 2nd ed. Torrance, California: Microcosm, Inc, 1992.

665-691. [9.6] Strom, Steven R. “International Launch Site Guide.” 2nd ed. El Segundo, California: The

Aerospace P, 2005. 117-128.


Final Report 105

10.0 – Propulsion System In spacecraft propulsion there are four primary methods used to maneuver the spacecraft: Solid, Liquid, Cold Gas, and Electric. A solid propellant is typically a fuel similar in composition to a pencil eraser that when provided with enough energy will burn until the fuel has been extinguished. This type of system is what most rockets use for their first stage. Well known examples of solid rockets are the white SRB’s located on the sides of the Space Shuttle. Liquid propellant systems typically have fuel and oxidizer tanks as well as pumps or pressurization systems to move the liquids through the system. The liquids are injected into a chamber where energy is added, normally in the form of a spark of flame, and combustion occurs. This type of system is used for the Space Shuttle Main Engines which are fueled via the large orange fuel tank seen between the SRB’s and the shuttle. Cold Gas systems are typically used for very low thrust applications such as stabilizing a spacecraft or pointing the craft in a new direction. In essence a cold gas system is the emptying of a gas or liquid into space. These types of systems are typically seen on small nanosats and some commercial satellites. Electric Propulsion engines operate on the principle of accelerating a cold gas to high velocities before exiting the spacecraft. This can be done via numerous methods, but the most popular include heaters, spark plugs, and particle accelerators. 10.1 - Scope and Purpose The propulsion system is divided into attitude control, altitude maintenance, and auxiliary propulsion systems. This section will include, requirements, performance, hardware, configuration, operation, and design evolution. The purpose of the Odyssey propulsion system is three-fold: the attitude control subsystem assists the Odyssey attitude control system (ACS) in maintaining the stability and control of the vehicle. The attitude control thrusters are used primarily to reset the ACS reaction wheels. They can also be used for rapid changes in spacecraft orientation and can provide emergency attitude control for the vehicle in the event of an ACS failure. The altitude maintenance subsystem maintains the altitude of the spacecraft’s orbit using an electric propulsion thruster to counteract the atmospheric drag forces on the vehicle. There are also auxiliary propulsion systems for modules, segments, and the super-module. These propulsion systems transport the module or segment from the launch vehicle to Odyssey. After docking, these units are detached and deorbited.


Final Report 106

The altitude maintenance and attitude control subsystems utilize hydrazine for fuel, Helium for pressurization, and share common fuel and pressurization tanks. These propulsion subsystems can be refueled and repressurized by the Orbital Express spacecraft. Further propulsion modules can be added due to the integration of fuel connections with the segment docking ports. A large Helium reserve is carried to allow for the refueling of future free flying modules.

10.2 - Tasks, Functions, Requirements, and Design Methodology The Odyssey propulsion system is designed to fulfill two tier one requirements; that the propulsion system provide for altitude maintenance and attitude control, and that Odyssey must be modular. Section 2.3.1 lists the system requirements pertinent to the propulsion subsystems. Altitude maintenance, per requirement 1.2, is achieved using a single 1 kW hydrazine arcjet located at the aft end of the propulsion segment to provide a ΔV increase of 1m/s/yr. The electric propulsion arcjet was chosen over a chemical propulsion alternative for several reasons. The arcjet offers a higher efficiency than chemical hydrazine monopropellant thrusters, and there is an overall weight savings due to a lower propellant mass requirement. Furthermore, the arcjet is fired for long periods at a very low thrust. This allows the spacecraft to continue its imaging duties while the spacecraft is being re-boosted. Attitude control is achieved using 12 hydrazine monoprop thrusters located in four clusters of three, with two clusters at each end of the spacecraft. These thrusters are used to unload the ACS reaction wheels once every four days per requirement 1.4. The system carries sufficient propellant to rotate the spacecraft at a faster velocity than the reaction wheels can alone as specified in requirement 1.5. In an emergency the attitude control thrusters can be used in place of the reaction wheels to stabilize and control the spacecraft, though far field imaging capability is lost. The attitude control thrusters can also be used for altitude maintenance in the event of an arcjet failure. However, far field imaging capability will be lost during the re-boosting maneuver. The altitude maintenance and attitude control subsystems utilize common propellant storage and plumbing per requirement 1.6. Three 55.2 MPa Helium tanks are used to pressurize three 5.52 MPa hydrazine tanks. Helium was chosen for pressurization because it can be stored as a gas and at low temperatures requiring no cryogenic provisions. Hydrazine was chosen as a fuel because it is the most common monopropellant currently in use and is frequently used for arcjets. Both the Helium and hydrazine tanks can be refueled by the Orbital Express spacecraft per requirement 1.7.1. Odyssey is designed to be refueled annually but carries a one year emergency reserve of hydrazine. The spacecraft carries four times the requisite amount of Helium needed for annual propulsion operations. This additional Helium supply can be used to refuel free-flying modules per requirement 1.8. In addition, enough propellant and Helium is carried to safely deorbit the spacecraft at any time per requirement 1.3. The Odyssey plumbing configuration allows for dual redundancy in leak isolation as demanded by requirement 1.9. Modules, segments, and the super-module are launched with auxiliary propulsion units. These units are used to rendezvous and dock with Odyssey after the launch vehicle delivers the module or segment to orbit. The segment and super-module propulsion units utilize eight small


Final Report 107

hydrazine monopropellant thrusters for attitude control and a single large hydrazine thruster for altitude changes. The units are detached and deorbited after use per requirements 2.6 and 2.7. The module propulsion units utilize eight Helium cold-gas thrusters for attitude control and altitude changes. These units can detach from the new module after delivering it to Odyssey and reattach to an old module currently docked with the spacecraft. This allows for the repositioning of modules and the deorbit of old modules per requirement 2.5. 10.3 - Choices - Trade Studies, Decision Justification, and Design Drivers The primary design driver for the Odyssey propulsion systems is requirement 1.7 that the spacecraft be refueled annually. This directly impacted the configuration of both the altitude maintenance and attitude control thruster subsystems. Additional design drivers included the requirement for the spacecraft to be refuelable by Orbital Express (1.7.1) and that the two subsystems share common fuel storage (1.6). The first task in the propulsion design process was to select a propulsion method for altitude maintenance. Electric, chemical monopropellant, chemical bipropellant, and solid propellant thrusters were considered. Solid propellant thrusters were deemed unsatisfactory because they are one time use, not easily refueled/reloaded, and have fixed burn durations. Chemical bipropellants were also deemed unsatisfactory due to the complexity of having to store and refuel two systems. A detailed trade study was conducted to select between electric arcjet and chemical monopropellant thrusters. Before the trade study could be completed it was necessary to select a fuel and pressurizing system for the attitude control system. This decision was facilitated by the addition of requirement 1.7.1 mandating that the spacecraft be refuelable by Orbital express. Hydrazine was chosen as the propellant for the Odyssey because it is a monopropellant, meaning that it can be decomposed into a fuel and oxidizer; eliminating the need for a second propellant and reducing the overall system complexity. Hydrazine monopropellant thrusters are the most common form of satellite maneuvering thrusters. They are much simpler than bi-propellant systems while maintaining a high efficiency (Isp 210-230s). Hydrazine is the only monopropellant that has been demonstrated for use in arcjet electric propulsion (Isp 400-500). Most importantly, it is the only fuel currently compatible with the Orbital Express refueling system.[10.1] Helium was chosen as the pressurizing gas for the fuel system because it is easily stored as a compressed gas in space. Due to its low molecular weight, the use of Helium dictates welded plumbing, special valves, and expensive face-seal connectors to prevent leakage. Alternative pressurizing gasses like nitrogen and argon must be stored as liquids, requiring heavy and costly cryogenic connectors and valves along with the added complexity of strict thermal control. The Orbital Express spacecraft is currently configured to deliver Helium[10.2].

The altitude maintenance propulsion trade study was conducted using several assumptions. The spacecraft mass before propellant and propulsion hardware addition was estimated at 9,500 kg. An annual ΔV requirement of 1 m/s for three years was used, allowing a safety window of two


Final Report 108

years between refuelings. A 900 N hydrazine monoprop was selected as one candidate with an Isp of 220 seconds and a mass (including plumbing, shielding, etc) of 200 kg. A 1 kW arcjet with an Isp of 420 seconds and a mass of 7kg was selected for comparison. A mass penalty of 220 kg (1000 W at 4.5 W/kg) was added to include the power mass required to operate the electric thruster. Using the basic rocket equation (equation 10), the systems were plotted against one another with overall system mass for the requirement of 3m/s of ΔV being the deciding metric. The results are shown in Figure 10.1.

⎟⎠

⎞⎜⎝

⎛ −=Δ−

spoIgV

spacecraftpropellant eMM 1 Equation 10.1

The chemical monoprop thruster system was found to be only 5 kg lighter than the arcjet system, however, it uses 24 kg more propellant than the arcjet system. Since the system masses are nearly identical, it was determined that the next most important performance metric was fuel consumption over the lifetime of the vehicle. Further examination revealed that the chemical system would use 162 kg more propellant than the electric system would over the course of the propulsion segment’s 20 year life span. This led to the selection of the 1 kW hydrazine arcjet as the altitude maintenance thruster.

Altitude Maintenance System and Fuel Mass vs. Vehicle Delta V

200210220230240250260270280290300

0 2 4 6 8 10

Delta V (m/s)

Engi

ne a

nd P

rope

llant

Mas

s (k

g)

ArcjetChemicalLinear (Arcjet)Linear (Chemical)

Figure 10.1 – Combined Altitude Maintenance System and Fuel Mass vs. Vehicle Delta V

Selection of the attitude control thrusters was considerably easier. It was determined by the propulsion team that the attitude control thrusters should be able to rotate the spacecraft faster than the ACS reaction wheels. This would allow the spacecraft to rapidly change orientation for docking and emergency maneuvers. Pulsing hydrazine monoprop thrusters were chosen for this role. Arcjets were infeasible because of low thrust performance and high power requirements.


Final Report 109

The number of hydrazine storage tanks and Helium pressurization tanks was calculated with redundancy in mind. Three hydrazine tanks allow for a single tank failure without affecting spacecraft operation, and a two tank failure without sacrificing any mission critical spacecraft capability. Three Helium tanks allow for a double tank failure without affecting spacecraft operation. 10.4 - Design Evolution The design of the Odyssey propulsion segment has changed considerably from initial concepts. Some of the design concepts are shown in Figure 10.2. One of the initial concepts favored by the design team called for a geodesic shaped spacecraft with an arcjet protruding from one end. This design featured attitude control thrusters grouped into removable thruster modules.

Figure 10.2

Evolution of the Odyssey Propulsion Segment. Counterclockwise from top-left: the geodesic concept, dual propulsion system for plane changes and removable tanks, dual arcjets with 16 ACS thrusters, final

configuration with single arcjet and 6 ACS thrusters As the design progressed the geodesic concept was abandoned in favor of an octagonal cylinder arrangement. This design featured arcjets for altitude maintenance, a large chemical monopropellant thruster for plane change maneuvers, and thruster banks mounted on the segment docking collars for attitude control. The design also featured external fuel tanks that could be jettisoned, allowing for a “tug” to deliver replacements. The need for to conduct plane change maneuvers was eliminated with the selection of a polar orbit. The decision to use Orbital Express for refueling eliminated the need for external, removable fuel tanks. A new design emerged incorporating two arcjets for altitude maintenance


Final Report 110

and 16 hydrazine monopropellant thrusters for attitude control. An additional 16 attitude control thrusters were attached to another propulsion segment at the opposite end of the spacecraft. The number of attitude control thrusters was later reduced to 12 (24 total) to eliminate plume impingement on the spacecraft optics. This number was reduced further to eight (16 total) with the introduction an off-axis pulsing scheme whereby the thrusters were mounted in pairs and angled 45° off of two principle axis. This configuration proved to be too inefficient. The final configuration utilizes four clusters of three thrusters, with two clusters at either end of the Odyssey for a total of 12. The second arcjet, originally added for redundancy, was omitted after it was found that the attitude control thrusters could be used for altitude maintenance in an emergency.

10.5 - Final Propulsion Concept Overview Figure 10.3 shows the Odyssey propulsion system with its various subsystems highlighted.

Figure 10.3 – The Odyssey Propulsion Subsystems

The altitude maintenance subsystem consists of a single 1 kW hydrazine arcjet housed in the propulsion segment. This thruster is used to counteract the atmospheric drag forces on the spacecraft and maintain orbit altitude. This subsystem is explained in section 10.6.1. The attitude control thruster subsystem assists the spacecraft ACS system with maintaining control of the Odyssey’s orientation. The system consists of twelve 45 N thrust hydrazine


Final Report 111

monopropellant thrusters arranged in clusters of three at either end of the spacecraft. This subsystem is explained in Section 10.6.2. The fuel distribution subsystem consists of three 5.52 MPa hydrazine storage tanks, three 55.2 MPa Helium pressurization tanks, and all of the fuel and pressurization plumbing. The subsystem provides propellant to the altitude maintenance and attitude control subsystems. It also provides Helium for future applications. The subsystem is refuelable via Orbital Express through a docking port on the propulsion segment. This subsystem is explored in Section 10.6.3. The module propulsion subsystem provides attitude control and primary propulsion for new modules being launched for rendezvous and docking with Odyssey. It can also be used to reposition or remove existing modules. The subsystem utilizes eight Helium cold gas thrusters for propulsion and is explored in Section 10.6.4. The super-module propulsion subsystem provides attitude control and primary propulsion for the propulsion system for rendezvous and docking with Odyssey. The subsystem utilizes eight hydrazine monoprop thrusters for attitude control and a single large hydrazine monoprop thruster for orbital changes and is explored in Section 10.6.5. The segment propulsion subsystem provides attitude control and primary propulsion for new segments being launched for rendezvous and docking with Odyssey and is a derivative of the super-module propulsion subsystem. The subsystem utilizes eight hydrazine monoprop thrusters for attitude control and a single large hydrazine monoprop thruster for orbital changes and is explored in Section 10.6.6. 10.6 – Subsystem Design, Specification, and Operation The following sections give specifics on the configurations and uses of the propulsion subsystems. Detailed calculations and numerical justifications for each subsystem can be found in Appendix A.10.1 through A.10.6.

10.6.1 – Altitude Maintenance Subsystem The altitude maintenance subsystem (AMS) is designed to maintain Odyssey’s operating altitude of 792 km. Atmospheric drag on the Odyssey’s spaceframe results in a one m/s annual decrease in spacecraft velocity. The purpose of the AMS is to offset this decrease in velocity by using a hydrazine arcjet to reboost Odyssey to the nominal velocity of 706.96 m/s. The AMS uses a 1 kW hydrazine decomposition arcjet for propulsion. An arcjet is a resistojet electric propulsion unit which uses an electrical arc between an anode and a cathode to ionize propellant. The ionized propellant is then accelerated and expelled through a nozzle. The arc temperature within the thruster ranges between 10,000 and 20,000, yet the arcjet is inherently self cooling. This is due to the arc being unsteady, allowing the flow of unheated propellant to cool the thruster.[10.3]


Final Report 112

The arcjet specified for Odyssey is a custom design 10 cm in diameter and 0.5 m in length. The performance characteristics for thruster were derived using the performance data for constricted hydrazine arcjets found on page 558 of Space Propulsion Analysis and Design.[10.4] The arcjet produces 0.17 N of thrust at a hydrazine flow rate of 41±0.5 mg/s of propellant. The thruster has a specific impulse of 420 seconds and an efficiency of 33%. Including thermal controls, mounting, electrical conditioning equipment, and instrumentation the thruster weighs 7 kg. The reboosting burn consumes 2.42 kg of propellant and lasts 18 hours. This can be accomplished through several short burns or one long duration burn. The maneuver is completed during Odyssey’s annual maintenance systems check when no imaging is being performed. However, the low thrust of the arcjet allows for its firing during imaging sessions; provided that the spacecraft is properly oriented with the thruster firing along the spacecraft’s orbital path. The thruster draws one kilo-watt of power during operation. Between firings this power supply can be utilized by other spacecraft systems. The cathode of the arcjet decays with use over time. Current station keeping arcjets have a lifespan of 10-15 years. Based on the current level of technology, the AMS lifetime is conservatively estimated to be 15 years. However, current research efforts should yield a longer lasting cathode design in time to incorporate it into the Odyssey thruster.

10.6.2 – Attitude Control Thruster Subsystem The attitude control thruster subsystem (ACTS) works in conjunction with the Odyssey attitude control system to control the spacecraft orientation. The ACS uses reaction wheels to maintain the pointing accuracies required for imaging. These reaction wheels reach peak angular momentum about every 4 days. When that happens, the wheels are decelerated to unload their angular momentum. Thrusters are used during the unloading procedure to maintain the orientation of the vehicle.[10.3] The ACTS system consists of twelve 45 N hydrazine monopropellant decomposition thrusters. These thrusters decompose hydrazine across a heated Iridium catalyst bed into gaseous ammonia and nitrogen.[10.3] They are arranged in sets of three as shown in Figure 10.4. The thrusters are arranged along the principle axis of the spacecraft but are canted outboard 45°. This canting was done for three reasons: each thruster can affect motion along two axes, the exhaust plumes are directed away from the optics, and canting the thrusters allows for fine control of the spacecraft as the center of mass varies in position. The ACTS thrusters operate in a pulsed mode with a specific impulse of 220 seconds. Figure 10.5 shows the naming convention for maneuvers. The thrusters are fired in pairs for pitch maneuvers and in fours for roll and yaw maneuvers. The frequency and duration of the thruster pulses can be varied. Thruster firing is directed by the Odyssey computer system, while a controller takes measurements from the ACS gyros and compensates for drift by varying the frequency and duration of pulses.


Final Report 113

Figure 10.4 – ACTS Thruster Cluster Showing 45o Cant

Figure 10.5 – Odyssey Attitude Control Maneuvers

A propellant mass of 16.35 kg is required annually to unload the pitch control reaction wheel. 2.55 kg of propellant are required annually to unload each reaction wheel. A propellant mass of 16.35 kg is allotted to the ACTS to provide Odyssey with the capability to rotate at up to 3.14 rad/sec on the pitch and yaw axes and 1.13 rad/sec on the roll axis. This allows the spacecraft to reorient at a faster speed than can be achieved using reaction wheels alone. The capability can be utilized for rapid retargeting of the optics, or for emergency maneuvering. The ACTS system can also be used for altitude maintenance in the event of a failure with the AMS and for complete attitude control in the advent of an ACS failure. While the ACTS can safely stabilize, control, and reboost the Odyssey, it cannot maintain the pointing accuracy required for far-field imaging. The thruster positioning allows for multiple failures without loss of maneuvering capability, though firing schemes will become more complicated to compensate for failed thrusters.


Final Report 114

10.6.3 – Fuel Distribution Subsystem The fuel distribution subsystem (FDS) supplies fuel to the AMS and ACTS. It consists of three 5.52 MPa hydrazine tanks, three 55.2 MPa Helium tanks, plumbing, valves, and instrumentation. The hydrazine and Helium tanks are refuelable via orbital express. Figure 10.6 shows the arrangement of the FDS hardware within the propulsion segment. A 998.77 kg mass of hydrazine is stored among three 0.957 m diameter spherical titanium fuel tanks. The tanks, produced by ATK Space Operations, have elastomer bladders and weigh 35.5 kg empty. A single tank can fail without significantly effecting vehicle operations. The spacecraft can survive on a single tank, but will not have enough propellant to deorbit.

Figure 10.6 – Arrangement of FDS Tanks Inside Propulsion Segment

A 107.30 kg mass of Helium are stored among three 0.51 m diameter spherical titanium pressurization tanks. The tanks, also produced by ATK Space Operations, weigh 41.89 kg each empty. Each tank contains enough Helium to fully pressurize all three fuel tanks with a 25% emergency reserve. This allows the spacecraft to operate without limitation on a single tank Figure 10.7 shows the FDS plumbing and instrumentation diagram (P&ID). The fuel lines are 0.5” x .065” stainless steel tube. The Helium valve actuation lines, and vent lines are 0.25” x .035” stainless steel tube. The Helium pressurization lines are 0.5” x .083” stainless steel tube. Valves are supplied by Moog and Marotta. Seal-Loc tube fittings are used for all connections and are supplied by Parker. The Helium lines are welded to reduce the risks of leaks developing over time.


Final Report 115

The FDS plumbing is of a single fault tolerant configuration. The system can tolerate at least one valve leak in any line without losing any operating capability. The system can tolerate at least a double failure in any line without significant loss of capability.

Figure 10.7 – FDS Plumbing and Instrumentation Diagram The hydrazine and Helium tanks are refilled annually by Orbital Express, shown in Figure 10.7. Orbital Express is a DARPA program to demonstrate autonomous refueling and repair operations. The current version of the refueling spacecraft is called ASTRO.[10.5] A docking connection for the Orbital Express spacecraft is located on the back of the propulsion segment. The Orbital Express spacecraft connects with the Odyssey using Vacco quick-disconnect fueling ports.[10.1] The pressure in the Odyssey hydrazine tanks is then vented allowing the tanks to be refilled. The Helium tanks are repressurized by opening the Helium connections between the two spacecraft.


Final Report 116

Figure 10.8 – Orbital Express ASTRO Vehicle (left) Servicing

a communications satellite (right). [10.6]

10.6.4 – Module Propulsion Subsystem The Module Propulsion System (MPS) is designed to bring new modules to Odyssey. Loaded onboard a Falcon 5 launch vehicle, each system is capable of carrying an individual 1m long module. The system will maneuver the module from payload fairing separation through docking with Odyssey. After docking, the MPS will disconnect from the module, maneuver to and dock with an old module, then proceed to deorbit the old module. The MPS system was designed to be a lightweight, inexpensive, “throw-away” system. To accomplish this, the propulsion system was designed to be compact and simple. Numerous methods were considered, including solid propellants, hybrid systems, monopropellant thrusters, bipropellant thrusters, and electric propulsion systems. A Helium cold gas thruster system was chosen to be the most beneficial. The Helium cold gas thruster is able to provide the required level of thrust and has the advantage of not needing any thermal system to remain in liquid or gaseous form. The MPS carries enough Helium to complete its entire mission through the deorbit of a module to an altitude of 80km; where the Earth’s gravity will rapidly finish deorbiting the MPS and module. This is below what is considered “space” and the MPS and module will burn up well before the 80km end-of-fuel limit. The MPS will deploy three antennas operating at frequencies of 2.14, 2.16, and 2.18 MHz to properly orient the spacecraft after exiting the payload fairing. Odyssey will locate each antenna and determine the orientation of the spacecraft. Odyssey will then direct the MPS from its orbit insertion through docking using its ACS computer and LADAR. The MPS will process commands from Odyssey using a small 166 MHz processor.


Final Report 117

Figure 10.9 - MPS with module attached

The total power consumed by the MPS is 60W and is supplied by onboard batteries. The batteries can sustain the MPS for 8-10hrs, depending on power consumption. After this time the MPS will reorient so that its solar array is pointed towards the sun; trickle charging the batteries. Approximately 1.5 hrs of direct solar charging is needed to fully charge the batteries. The MPS system is constructed of titanium alloy to ensure structural rigidity and lightweight. The titanium alloy reduced the total structural mass of the MPS to 70 kg. The total system mass is approximately 1000kg including 570 kg fuel tank, 270 kg of fuel, 15 kg of batteries, and 100 kg of additional hardware. The MPS system can maneuver a module of up to 1000kg 12km to dock with Odyssey and then deorbit an old module. The system is not designed to be a permanent fixture on Odyssey, but is not limited to short duration spaceflight. The goal for the MPS is allow for it to be docked to the Odyssey for up to 6 months before performing a deorbit burn. This will allow for new modules to be launched without having to wait for an existing module to expire. The overall material cost for the MPS is approximately $150,000 with the total MPS cost will be less than $250,000 for a single unit after assembly and testing. As additional MPS units are produced the overall cost will decrease leading to additional customer savings. In summary, the MPS is the workhorse that allows for the replacement and addition of new modules to Odyssey. This critical component makes the system modular and upgradeable, performing the complex task of docking and undocking modules.


Final Report 118

10.6.5 – Super-module Propulsion Subsystem A propulsion system is needed to dock the super-module containing the space viewing telescope to Odyssey. This system was designed as a scaled up version of the MPS. The structural calculations reveal that a linear volumetric scale will result in a safety factor of four which, although in excess of the required safety factor of two, reduces computational cost and allows for a quick design of the system. The revised structural model, shown in Figure 10.10, has a SF of four and is designed to withstand a 7g launch loading. The additional bottom plate was added to reduce potential torsion strains on the super-module. The size of super-module dictates a propulsion solution more capable than a Helium cold gas thruster system. A hydrazine monopropellant thruster with Helium pressurization was chosen for this system for compatibility with other systems used on The Odyssey Project. The overall system will also have a primary thruster, and will retain the computer, antennas, battery, and solar array used for the module propulsion segment. Because only one super-module is to be used the design has not been the primary focus to this point. The conceptual design and preliminary first order analysis completed, however a detailed design has not yet been completed. The system will need to transport approximately 3000 kg, 10 km to Odyssey and then detach and deorbit. The final design of this system will occur in sufficient time to meet the launch and construction schedule.

Figure 10.10 - Super-module Structural Model

10.6.6 – Segment Propulsion Subsystem


Final Report 119

The segment propulsion system (SPS) is a hybrid design between the primary propulsion system for Odyssey and the super-module propulsion system. It is shown in Figure 10.11. Because of the potentially large mass of a segment and the short launch to docking time an electric propulsion system would be impractical. A hydrazine thruster system was chosen similar to the one used for super-module propulsion. This system employs eight directional thrusters and one primary thruster to maneuver the new segment into position.

Figure 10.11 - Segment Propulsion Subsystem

Structurally the system is similar to the primary propulsion system and will be cylindrical in shape, and smaller in size. The segment propulsion system will connect to the new segment via the same core docking mechanism used for the primary propulsion system. This means that the SPS can potentially be refueled via its core connection and reused for other applications. The exterior of the cylinder will be covered with solar panels to charge the batteries used to power the computer, valves, and additional hardware. Overall the system will be capable of maneuvering a 5000 kg segment 10 km, docking with Odyssey, and maintaining the capability to deorbit the SPS. This is a low cost system and designed to be disposed after use, but there is potential for additional re-use due to refueling capabilities. 10.7 – Propellant Budget and Refueling Schedule Figure 10.12 breaks down the FDS hydrazine supply. The Odyssey is designed to be refueled annually. The system carries enough Helium for up to four years of operation, but only has a one year reserve of hydrazine. This reserve was included to account for the possibility of a launch failure of the Orbital Express refueling vehicle.


Final Report 120

Table 10.1 – Hydrazine Supply Breakdown

Hydrazine Mass (kg)

Altitude maintenance 2.42Reaction wheel unloading 7.66

Attitude control maneuvering 264.57Total annual requirement 274.65

Deorbit requirement 464.82Total requirement 739.47

FDS Capacity 998.78

Reserve 274.65

The 276.31 kg propellant reserve can also be utilized during periods of frequent maneuvering, though nominally this reserve would remain untouched. In an extreme emergency the deorbit allotment of 467.04 kg can be tapped to provide propellant for up to 1.7 years of additional operation. 10.8 – System Life Expectancy The shortest time limiting factor on the life of the AMS, ACTS, and FDS is the life of the AMS arcjet. Should the arcjet fail, ACTS can be used for altitude maintenance, but at lower fuel efficiency. The ACTS thruster Isp will degrade from 230 to 210 seconds over the 20 year lifespan of the system. However, this will not reduce performance; merely accelerate the refueling schedule by a few months. The FDS system features a high level of redundancy so that it can operate over the 20 year lifespan of the vehicle. It is important to note that the propulsion segment is removable and can be replaced in the event of an extreme failure. The ACTS thrusters on the optics segment may no longer be used once the experiment module is added, as a new bank of thrusters could also be added. The module, segment, and super-module propulsion systems are all considered disposable. 10.9 – Cost Estimation The cost for the Odyssey propulsion system is estimated at $4,420,000. The breakdown of the component costs can be found in Appendix 10. The module propulsion system is estimated at $250,000 per unit. Segment propulsion modules are estimated to cost $2,000,000 each and the super-module propulsion unit will cost $4,000,000.


Final Report 121

10.10 – Summary The design challenges of refuelability, reliability, control, and delivery were creatively satisfied by the design of the propulsion system. Altitude maintenance is provided by electric propulsion, minimizing imaging disturbance. The attitude control system satisfies unloading requirements and also enables rapid maneuvering and emergency control. Both systems utilize a common propellant, pressurizing gas, and supply system to facilitate refueling via the Orbital Express system. The propulsion system is single fault tolerant and can accept at least two connected failures without endangering the spacecraft. Finally, auxiliary propulsion units provide a means for modules and segments to be transported from their launch vehicles and docked with the Odyssey. The investigation of the propulsion design challenges has led to several lessons. The evolution of a propulsion design is married to the tumultuous definition of the spacecraft mass. Propellant requirement is a function of spacecraft dry mass. Yet any increase in propellant mass results in an increase in propulsion system mass which further increases the dry mass of the system. These increases filter down to affect thrust requirements, response times, and geometric constraints. The cyclical nature of the issue made getting the actual propulsion system dry mass to converge with the estimated value one of the most difficult tasks of the analysis. Defining the propulsion system was facilitated by the Orbital Express refueling requirement, as it mandated hydrazine based propulsion units. However, there are other fuels, including monomethyl hydrazine (MMH) and Aerojet’s MMH hydrazine blends, which offer higher levels of performance but are not yet compatible with Orbital Express. Furthermore, technical setbacks within the Orbital Express system since September have placed the future of the program in doubt. Jetisonable fuel modules, originally defined for use, were rejected due to rendezvous and docking issues. Yet the development of the module propulsion system now makes this option feasible. In hindsight jetisonable tanks may have been a better solution to the refueling issue as they allow for greater flexibility in propellant selection. Future effort should be dedicated to developing iterative software for tabulating spacecraft mass. The Excel bookkeeping method utilized for the design was a linear system with propulsion analysis as an end point. As a result, any changes in propulsion system dry weight (i.e. tank masses) had to be adjusted at the system level and recalculated. Cyclic limitations made it impossible to use the excel solver functions to expedite convergence. An iterative Matlab mass calculator would have been much more difficult to implement, but would have allowed for higher accuracy in the mass predictions, and thus propulsion, analysis.


Final Report 122

10.11 – References [10.1] Dornheim, Michael A. "Orbital Express to Test Full Autonomy for on-Orbit Service."

Aviation Week 06 Apr. 2006. 2 Dec. 2006 <http://www.aviationnow.com/avnow/news/channel_space_story.jsp?id=news/aw060506p1.xml>.

[10.2] Sackheim, Robert L. A View of Future NASA/MSFC Propulsion and Space

Transportation Activities. Alabama/Mississippi Section Annual AIAA Meeting, 4 May 2006, NASA MSFC. 2 Dec. 2006 <www.aiaa.org/Participate/Uploads/Prop%20and%20Space%20Trans%20-%20Sackheim.pdf>.

[10.3] Sutton, George P., and Oscar Biblarz. Rocket Propulsion Elements. 7th ed. New York:

John Wiley & Sons, 2001. [10.4] Humble, Ronald, Gary Henry, and Wiley Larson. Space Propulsion Analysis and Design.

McGraw-Hill, 1995. [10.5] Kennedy, Lt Col Fred. "Orbital Express Space Operations Architecture." DARPA Tactical

Technology Office. 8 Aug. 06. USAF. 2 Dec. 2006 <http://www.darpa.mil/tto/programs/oe.htm>.

[10.6] "Orbital Express Space Operations Architecture." Tactical Technology Office. 8 Aug.

2006. DARPA. 11 Nov. 2006 <http://www.darpa.mil/tto/programs/oe.htm>.


Final Report 123

11.0 – Power System The power system is what allows the operation of the spacecraft to happen. It provides the necessary energy needed for day to day operation. Power systems for spacecraft commonly employ technologies like solar cells, batteries, electrical fuse protection systems, etc. All of these technologies, when combined in the proper manner, allows for a system that is concrete in nature and provides the proper solution when considering the power problem for the Odyssey. 11.1 – Scope and Purpose This section will cover the design of the Odyssey’s power system. The power system is divided into three main sections: generation, storage, and distribution and control. This section will explain how the power system for the Odyssey came to fruition. Functions, trades, and the process followed to complete the design will be discussed. The means of generation picked for the Odyssey would be some form of a solar array. Solar arrays are heavily used in the space industry and in our case would be determined to be the best solution to our problem. Things considered for this problem would be the power demand and availability of space on the Odyssey for a power generation system. Storage was accomplished with Nickel Hydrogen batteries. Presently, there are two choices for batteries which are commonly used by satellite systems and in our case, the Nickel Hydrogen batteries proved to be the best choice. Distribution and control is the section of the power system that covers all remaining aspects. The route chosen fulfill the needs of distribution and control would be a peak power tracking system. This would govern the control of energy from all inputs and outputs as a modular system would desire this. The power system of a satellite is essentially the heart of the operation. It provides the lifeblood necessary to the other systems that allows for their operation. Without a successful and efficient power system, our project would be hindered at many levels. The objective of this section will then be to provide insight at all levels behind the creation and operation of the Odyssey’s power system. 11.2 - Tasks, Functions, Requirements, and Design Methodology The power system for the Odyssey was required to meet many tier requirements ranging from levels of tier 1 to tier 3. Each one of the requirements was based around the three main sections of design: generation, storage, and distribution and control. The tier requirements are as follows: 12-The power system shall provide adequate power generation, 12.1-The power system shall provide adequate storage of power, 12.2-The power system shall offer adequate means to regulate and distribute power, 12.3-The power system shall provide a memory keep alive system, 12.2.1-The power system shall provide fuse protection for equipment, 12.2.2-The power system


Final Report 124

shall provide emergency load shed. These tier requirements would provide a basis for which each section of the power system would be designed to. The first objective to overcome was to decide on a reliable means of energy generation. Upon first glance, there were many options to be explored. However, in the end, solar energy was the winning choice. Some key points that led the designer to this choice were space occupied on the Odyssey and lifetime. Although solar panels are commonly seen as rigid bodies attached to the spacecraft, it was the Hubble Space Telescope that brought about an option not yet seen in our design. The Hubble uses a cassette style design which rolls out a solar blanket. In this way, it can provide a large solar area for power generation without occupying a large amount of space at launch.[11.9.2] The next area of concern was the storage of power. An effective means would be needed as the system could have on the spot power demands that the solar array may not be able to handle and the potential for nighttime loads existed. There are two types of batteries commonly used in industry at present and they are Nickel Cadmium and Nickel Hydrogen. Nickel Hydrogen batteries were chosen because they offer more discharge capability at a higher efficiency for the same amount of batteries as compared to Nickel Cadmium. The advantage of this is to save weight and space.[11.9.3] Power distribution and control was another problem to be solved when considering the design of the spacecraft. The area of concern here was to ensure that the Odyssey was equipped with a proper means of routing and controlling the power needed for the other subsystems. This entails making sure the design allotted enough space to fit the desired equipment needed for such operations. The methodology used in the design of Odyssey’s power system was as follows. The tier requirements for the power system are straightforward. The restrictive part lies within the fact that the power system had to be designed to meet the requirements of the entire spacecraft as well. Those were things such as modularity, compatibility with orbit, lifetime, etc. These were major things that affected the design process of the Odyssey’s power system. 11.3 – Choices: Trade Studies and Decision Justification, Design Drivers In coming up with the final design for each section of the power system, many considerations had to be made before a solution was accepted. One of the design drivers for the power system was the end of life power requirement.[11.9.3] At the beginning of the mission, Odyssey’s subsystems will have power demands that need to be met. However, due to a solar panel’s degradation over time, this power demand needs to be scaled up so that the same demand at the end of the suggested life of the Odyssey can be met. This played a substantial role in governing the amount of surface area actually needed for the solar array. Another design driving factor was the ability to provide modularity for the power system. Given a set number of modules and a size of these said modules, the power system would have to be designed to fit to these aforementioned restrictions. It turned out that this would not be an issue for the design of the Odyssey’s power system.


Final Report 125

When considering how to reduce weight and cost, an effort was made to see which systems did not necessarily need to be on at all times. The trade behind doing this was a reduction in amount of surface area needed by the solar array to generate necessary power and a reduction in the number of batteries required to store this said power. This reduction would also have an effect on the distribution and control aspect as less fuse protection would be required at the bus bar. Through communication with the other subsystem teams, a scenario was conceived that eventually contributed to the solar array design which was selected. 11.4 – Design Evolution From beginning to end, the design of the power system went through various phases of change, due in part to the change in functional requirements which were demanded of the system. The initial design was one of a simple nature which still met all requirements. However, as time passed the requirements for the system changed which caused the satellite itself to undergo certain physical changes. The design of the power system had to match these changes to stay compatible with the system. The initial design phase saw the power system consisting of only a solar power unit with no batteries and a direct energy transfer to all other subsystems. This was the first choice, as the initial orbit for the Odyssey was to be a sun synchronous orbit, meaning the solar array would always be in the sun. As time passed, this orbit was deemed unworthy and thus the power system had to be modified to the new LEO orbit that the Odyssey would inhabit. Having moved to a LEO orbit would now designate the need for a storage system for the power, since there would be flight time where the Odyssey would be without sunlight. This addition was simple enough as the batteries were easily chosen to fit this addition to the system requirements. Distribution and control of the power would now have to be considered for change, since another path of travel was added to the route of power flow. This would add weight and cost, but not at unallowable levels. From this point, the evolution of the design did not vary greatly. All aspects of the system were now in place and only minor variations would be made to accommodate for a shift in power demand, or replacement of power modules. 11.5 – Solar Array Design The design of the solar array went through a few strong design phases, until the final choice was reached. The final choice was that of a flex foldout solar array. The flex foldout design consists of a housing which holds a solar blanket in a stowed fashion for launch. The blanket is rolled up in this housing until it is need where it is then unfurled via an extending mechanism that operates much in the same fashion as a scissor lift does. Originally, the flex foldout array was the only means of power generation. However, given the launch timeline of the Odyssey, the entire spacecraft would not be assembled all in one launch. This was one of the original tier requirements of the system. Therefore, the flex foldout array was resized to only produce for the


Final Report 126

initial launch of the bus segment and propulsion systems. Further down the launch timeline, the super-module would be launched and attached to the super-module would be the rigid solar array which would have the required surface area needed. The flex foldout array would then become and extra means of energy generation for the experiment segment and whatever systems may need extra power on top of what could be provided by the rigid array. The area needed to supply power for our spacecraft came to approximately 66 m2. The material chosen for the solar array was Gallium Arsenide, and of the materials possible, it offered long life and maximum power production. Solar cells, by nature, are not that efficient so it was necessary to search for a material that could provide the best possible output. To see a graphical representation of the power module please refer to Figure 11.1 below.

Figure 11.1 – Flex foldout array and rigid solar array

11.6 – Battery Design Because Odyssey is required to operate without sunlight, a battery system was necessary for operation of the other subsystems. The choice of batteries for this system was Nickel Hydrogen (NiH). The reason for this choice was because NiH batteries offer better output and longer life at that same amount of batteries as the other commonly used battery in industry Nickel Cadmium. An effort was made to reduce weight of the entire power system and the batteries had the potential to add a good amount of weight to the system. The number of batteries the Odyssey requires at initial launch is four. The weight of the battery system is approximately 93 kg.


Final Report 127

Given the geometrical design of the modules, the batteries can be fabricated in a manner which will occupy two of the three allotted modules for the power system. 11.7 – Bus Bar design The last of the three modules will have space for the distribution and control function of the power system. The bus bar will provide the interface between the generation and power systems to the rest of the subsystems of the Odyssey. The feature will incorporate the fuse protection, emergency load shed, and memory keep-alive functions of the power system. The bus bar will operate at a 28 volt capacity, which is common in industry. The bus apparatus of the power system will also feature a peak power tracker to offer fully regulated control of each of the power system’s features. The fuse protection and emergency load shed features offered are much like one’s typical fuse box in one’s home. This system will offer protection from power overload, or overdraw for the other subsystems on the Odyssey. This option is in place to prevent failure, or damage of other subsystems. Furthermore, the memory keep alive is in place to allow for control of the switching on and off certain subsystems that may not be required to be on at all times.[11.9.1] This is done in an effort to conserve power and increase the lifespan of the power system. 11.8 – Material Cost In terms of relative cost of the other subsystems, the power system is not the most expensive. However, the solar array’s can be quite costly. The estimated cost of the solar array is 18 million plus or minus a few hundred thousand. This is based on the cost of the actual cell, which is around $75.00 per cell with the cell being a 2 cm by 2 cm dimension in surface area. This cost accounts for both the flex foldout arrays which are initially launched with the bus segment and the rigid array launched with the super-module. The cost estimation also accounts for all of the extraneous features tied to the solar assemblies. These are things such as the solar array housing, the attachments and motors that control movement, etc. The batteries are not nearly as expensive as the as the solar array, but still appropriate for the amount needed. The estimated cost of the batteries is $7,000.00. This price is based on a cost per kilogram of Nickel Hydrogen, which was conceived from previous systems that have used such technologies. 11.9 – Summary It was shown in this text exactly how the Odyssey’s power system generates, stores, and distributes and controls electrical power. The design of the solar arrays was shown and the methodology behind the selection process was given. The same is true for both the batteries and bus bar features of the power system. An emphasis was made to show the importance of each


Final Report 128

system of the power system and more importantly why this power system was designed the way it was. Without a reliable and proven power system, the operation of the Odyssey spacecraft is unlikely. It was the hope that this section proved that such a power system was designed.


Final Report 129

11.10 – References [11.1] Larson, Wiley J. Wertz, James R. “Space Mission Analysis an Design.” 3rd Edition,

Microcosm Press, El Segundo California [11.2] “Hubble,” 17 September 2006, Encyclopedia Astronautica.

http://www.astronautix.com/craft/hst.htm. [11.3] Brown, Charles D. “Elements of Spacecraft Design,” Wren Software, Inc., AIAA, 2002.


Final Report 130

12.0 – Thermal Control The Thermal Control System (TCS) is responsible for maintaining the necessary temperatures of all the systems on the satellite. Every system has a range of temperatures they can successfully operate within. They also have a wider range for when the system is not operational. Some systems such as sensors, mirrors and cameras have temperature gradients maximums that are usually less than 0.5° C. This means the temperature difference from one side of the component to the other side cannot be more than 0.5° C. Thermal expansion of the structure will affect the mirror or sensor if the gradient is too large and will cause misalignment and pointing errors. 12.1 – Scope and Purpose

Achieving these temperature ranges requires the TCS to account for the heat input (the Sun, the Earth and the electronics) and the emissions from the surfaces and antenna. The Sun emits 1371 W/m2, the Earth emits 256 W/m2 and up to 550 W/m2 comes from the Sun but is reflected off the Earth’s atmosphere.[12.1] All three of these energy sources can fluctuate depending on the atmosphere of the Earth and the distance to the Sun. Thermal engineers must design a system that can handle a multitude of extreme environments. The systems available to dissipate the heat fall under two categories: passive and active. Passive systems do not require power. Examples include things such as insulation, isolation, multi-layered insulation (MLI), second surface mirrors and louvers. Active systems require power to operate and sometimes will involve moving parts, moving liquid and computer controls. Examples include things such as heaters, cryogenics, and heat transfer pipes. It is always a goal to use passive control when possible because it is cheaper and has less chance for failure. The Thermal Control section will present the development of the TCS design for the Odyssey Project. This begins with an explanation of the steps that lead to a completed design. The process used to develop the environmental model is explained, and then all temperature ranges and electrical consumption are accounted for. Next an explanation is presented which describes the components available for use. Then the trade studies performed will be examined to show the thought process that led to the choices of system components. Finally, the design for the modules and segments is presented. The end of the section concludes with a summary and cost breakdown.

12.2 – Tasks, Functions, Requirements and Design Methodology The task of the Thermal Control System (TCS) is to maintain the proper temperatures for all components and remove all waste heat. This is broken down into five Tier 2 requirements: 1) limit the energy transfer taking place within the propulsion segment 2) keep the propellant (hydrazine) above 2º Celsius or else it will freeze 3) warm the electronics and dissipate the heat in the module as needed 4) maintain a temperature gradient less than 0.1°C across the mirrors 5) remove heat from the core.


Final Report 131

The design process for TCS begins with creating a model of the environment the satellite will encounter. Modeling the environment requires looking at four different sources of energy: Solar radiation, albedo, planetary emissions and atmospheric drag. For this project there is no atmospheric drag and the only planetary load is from Earth. The model must have worst case hot and worse case cold temperatures. Another crucial aspect of the environment is the orientation of the satellite. This will determine which side is being impacted by the various environmental loads. All possible orientations should be investigated to determine the extreme hot and cold cases. The next step will determine the temperature range of the components and their electrical consumption. Electrical systems and moving parts operate the best within a range based on their physical properties. Some systems prefer to stay at one end of that range. For example solar panels have the largest range at ±100o C but they generate more energy at lower temperatures. The electrical consumption of the components is an important part of thermal control. All the energy coming in must go somewhere when the energy has done its job. Electrical systems reject this energy through heat and radiation. Antennae transmit some of their energy through radio waves. The energy consumption varies and knowledge of when the variations occur is important. For example, the Earth viewing telescope is not used in the shadow of the Earth and decreases the energy load on that module by 180 W. The electrical usage is added to the environmental load equation. The equation for each surface is balanced by selecting an IR emissivity (ε), solar absorbance (α), heater power and the temperature ranges achieved by this combination. This step will require many iterations and choices. The goals of this step are to minimize the heater power, minimize the difference between the extreme temperatures and keep the temperature within a safe range. This means the temperature, while in the Earth’s shadow, will be near the lower end and the Sun facing side near the upper range. There should be a safety margin between the temperatures achieved and the operational temperature range values in case a heater fails. The driver of this choice is the electrical usage of the components and the frequency of their use. The most difficult modules have electrical usage that decreases by 100 W or more when in shadow. The second step is to choose components for the internal thermal systems especially radiators, heaters and multilayered insulation (MLI). MLI minimizes heat transfer from one component to another within the spacecraft. It can also be used on the outside of the spacecraft instead of a radiator. MLI on the outer surface minimizes the environmental load to the closest values to the opposite of a black body (an object that absorbs all wavelengths of radiation). Heaters are placed within any internal structure to maintain their necessary temperatures The cost of the system is estimated after the heater sizes have been chosen. Thermal systems are usually 2 to 5% of the power and mass and less than 2 to5% of the total cost. [12.2] Once the system is built, it is tested in a vacuum chamber. The tests determine if the system can protect the satellite in the environment the satellite will encounter. The test cycles through the highs and lows multiple times and exposes an extreme scenario that might jeopardize the


Final Report 132

satellite. The internal components will also be tested to see if the heaters, heat pipes, refrigerators and cryogenics will handle the load they are designed for.

12.2.1 Space Environment Thermal control is driven by the environment of the spacecraft. The three environmental loads transfer heat through radiation but they are at different wavelengths. The solar load and albedo cover the visible spectrum and the short infrared radiation (IR) wavelengths (0.2 to 2.6 micrometers). The Earth emissions and the emissions from the spacecraft are at longer IR wavelengths (5 to 50 micrometers). These two sets of radiation wavelengths are illustrated below in Figure 12.1. The solar load is on the left and the Earth load is at the same wavelength as the room temperature curve. The dashed line shows the absorbance/emittance of a quartz mirror. The amount of energy absorbed or emitted at a certain wavelength is equal for a given object (the quartz mirror emits and absorbs 3% at 1 μm). However at different wavelengths, that value may be different (the quartz mirrors absorbs 65% at 10 μm and 3% at 1 μm). This is important because it allows for some surfaces to emit a significant amount of IR but not absorb so much of the solar load. The quartz mirror does not absorb much of the solar load or albedo and can emit a large amount of long-wavelength IR.

Figure 12.1 Solar and Room-Temperature-Body Spectral Distributions [12.3]

The environment model begins with an orbit or trajectory selection. Altitude and inclination are two parameters that must be known to make the environment model. These two parameters determine the orbit beta angle (β). The orbit beta angle is defined as the minimum angle between orbit plane and the solar vector and it is defined mathematically as

)cossin)sin(sin(cossin 1 RIRI SSS δδβ +Ω−Ω= − Equation 12.1 where δS is the declination of the Sun, RI is the orbit inclination, Ω is the right ascension of the ascending node (RAAN) and ΩS is right ascension of the Sun (RAS). The beta angle ranges from -90º to 90º. Figure 12.2 shows the orbit beta angle in three separate polar orbits. It is more difficult to imagine the beta angle when the inclination is not 90°. The beta angle is not constant because it is depends on the declination of the Sun, RAS and RAAN. For sun-synchronous

Quartz mirror radiator absorptance or emittance


Final Report 133

orbits the effect of the change in the RAS and RAAN do not affect the beta angle. Therefore sun-synchronous orbits fluctuate up to 25° while other orbits can fluctuate by 160°. These oscillations can occur more than 10 times for a non-sun-synchronous orbits and only twice for the sun-synchronous orbit.

Figure 12.2: Orbit Beta Angle [12.4]

The beta angle and the altitude combine to find the eclipse time for that orbit. Figure 12.3 shows the correlation between the altitude and the orbit beta angle. As the altitude increases the highest beat angle that is in an eclipse decreases. Table 12.1 lists all the parameters related to the orbit that are relevant to the thermal environment. Equation 12.2 determines the eclipse fraction where R is the Earth’s radius (6378 km) and h is the altitude. The absolute value of the eclipse fraction must be less than β*, which is the beta angle at which the eclipse begins. Both of these equations assume the Earth’s shadow to be cylindrical.

*

*2/12

1

0

cos)()2(cos

1801

ββ

βββ

≥=

<⎥⎦

⎤⎢⎣

⎡++

= −

if

ifhR

Rhhf E Equation 12.2

⎥⎦⎤

⎢⎣⎡

+= −

hRR1* sinβ Equation 12.3


Final Report 134

Figure 12.3: Eclipse Fractions [12.5]

Table 12.1: Orbital Parameters

Altitude (km) 791 Inclination (deg) 98.6

Beta angle range (deg) 19 to 27 Minimum beta angle with no eclipse (deg) 62.8

Eclipse fraction 30% Now that the orbit is defined, the environmental load can be determined. Modeling the solar load is the simplest. The energy load is at 1371 W/m2 ± 50 W/m2 depending on the eccentricity of the orbit (perihelion in early January receives the most while aphelion in early July receives the least). The Sun emits energy that arrives at the Earth perpendicular to the ecliptic (the plane of Earth’s orbit). If a surface is not normal to the solar vector, then the solar load is multiplied by cosine of the angle between the solar vector and surface normal. This angle is called the orbit angle (θ) and is shown below in Figure 12.4.

Figure 12.4: Orbit Angle [12.6]

Some of the Sun’s ray will reflect off of the Earth’s atmosphere and will impact the spacecraft. This energy is known as albedo and is calculated as a percent of the Sun’s rays that are reflected off of the Earth’s atmosphere back to space. Increases in latitude can lead to higher values because more energy is hitting ice and snow. Land will reflect more than water. Albedo at the zenith will reflect straight up while the poles will reflect away from the Sun. Albedo values can


Final Report 135

range from 0% at the terminator (the line between the shadow of the Earth and the illuminated bright side) to 40% over the zenith but the average falls between 25%-30%. There is no albedo in the shadow of the Earth. Albedo is inversely proportional to orbit altitude and can be ignored at altitudes past 30,000 km. The albedo value used in the model was selected from Figure 12.5. The albedo value in this graph depends on the orbit angle, the angle between the surface normal and local zenith (represented by ρ in Figure 12.4), altitude and the beta angle. The five surfaces facing the Earth have ρ values of 180º, 135º, and 90 º. These surfaces will encounter orbit angles from 0º to 40º. The average albedo values are 311 W/m2 for 180º surfaces, 217 W/m2 for 135º surfaces and 83 W/m2 for 90º surfaces.

Figure 12.5: Albedo Variation Due to the Surface Normal Angle [12.6]

The third type of environmental load is the Earth emission. Earth will release energy from the surface as long-wavelength IR. These rays will radiate in all directions. The Earth IR depends on the surface temperature. Thus it will be higher at equatorial latitudes than the upper latitudes where the surfaces are colder. Clouds can decrease the Earth radiation because they trap in the radiation. Earth IR is present even while in the Earth’s shadow, however the dark side will be colder. The typical value for Earth IR at the surface is 256 W/m2. Earth IR is inversely proportional to orbit altitude and can be ignored at altitudes past 30,000 km. Figure 12.6 shows the amount of Earth IR for various altitudes and ρ combinations. The Earth IR values that are used in this model are 166 W/m2 for 180º surfaces, 120 W/m2 for 135º surfaces and 50 W/m2 for 90º surfaces


Final Report 136

Figure 12.6: Earth IR Emission Variation Due to the Surface Normal Angle [12.7]

The final aspect of the space environment deals with the destructive properties of the molecules in the outer Earth orbit. Atomic oxygen (AO) is the most damaging in LEO [12.8]. AO is formed from molecular oxygen being split by UV photolysis. These atoms ram into satellites traveling at 8 km/s in LEO which imparts a large amount of energy. The chemical reaction between the valence electrons of the oxygen and the hydrocarbon polyimide surfaces (Kapton and Teflon) can form CO, CO2, and H2O. These products evaporate and expose inner surfaces. Compensation for AO is done by coating Kapton and Teflon with protective Aluminum coatings and orbiting at an altitude of 800 km and higher. Figure 12.7 shows the extreme concentrations of many particles including atomic oxygen.


Final Report 137

Figure 12.7: Atomic Oxygen Concentrations [12.9]

The temperature ranges for the Odyssey Project are given in Table 12.1. The module power usage is listed in Table 12.3. The two power columns list the power usage while in the sunlight and the Earth’s shadow. Three modules use different amounts because they are part of the EVT. The EVT is not in use in the shadow and therefore half of the computers are not being used.

Table 12.3: Odyssey Temperature Ranges [12.10] [12.11] Temperature Range(°C)Reaction wheels -19/49 NiH2 batteries -10/20 Rate sensing units -19/49 Computers -20/60 Fine guidance sensors -40/40 Solar Array Blankets -100/100


Final Report 138

Table 12.4: Module Power Usage

Module # Item Power in Sun (W)Power in shadow (W)1.1 RSU1 40 40 1.2 Computer 100 50 1.3 Antenna 5 5 1.4 Batteries 20 20 1.5 RWA1 200 200 2.1 RSU2 40 40

2.2 & 2.3 EVT 180 0 2.4 Empty 2.5 Batteries 20 20 3.1 RSU3 40 40 3.2 Computer 100 50 3.3 Antenna 5 5 3.4 Batteries 20 20 3.5 RWA2 200 200

Super Module FGS 70 70 12.3 – Thermal Components There are two categories of methods for thermal control: passive and active control. Passive control involves multi-layered insulation (MLI), radiators, conductive isolation, and louvers. Active control includes heaters, cryogenics, heat pipes, phase-changing material, air conditioners and deployable radiators. MLI is used to prevent radiation between spacecraft modules and the environmental load from space. MLI is made of 5 to 30 sheets of thin material (0.08 to 11 mil thick) separated by small conductive spacers (usually made of Dacron or Nomex). Typical materials that are used are Kapton, Teflon, Aluminum, Gold and polyester. The volume between blankets is evacuated to prevent conductive heat transfer. The pressure must be below one torr (0.0013 atm) for a noticeable decrease in conductive heat transfer from the air molecules. The sheets are loose enough not to tear but not too loose that they touch. The outer sheet is the thickest (0.5 to 11 mils) so it can protect the lower layers from micrometeoroids and thermal loads. The outer layer must also be resistant to surface degradation from charged particles, Ultraviolet (UV) radiation, Atomic Oxygen and contamination. The outer cover α/ε ratio is moderate (α = 0.4 to 0.55 and ε = 0.5 to 0.8). The middle layers are the thinnest (0.08 to 5 mils) because they are protected by the outer and inner cover. The have low emittance and are perforated to allow for ventilation of gases. The separators between layers need to have minimal contact with the layers, large thicknesses (6.5 mils) and have low thermal conductivity. The inner cover needs to be thicker (0.5 to 5 mils) to protect the inner layers. The inner layer needs to avoid electrical shorts and catching fire. Mylar is not used because of the possibility of flammability, and aluminum should not be used on the side facing the satellites because of electrical concerns. Very few varieties of insulation are made for an individual satellite. This reduces the manufacturing and design costs.


Final Report 139

Radiators and second surface mirrors are types of outer surfaces that are designed to radiate large amounts of energy. These surfaces use materials such as quartz mirrors, silvered and aluminized Teflon, optical solar reflectors and white paint to create a surface that emits large amounts of IR radiation. Typical α/ε ratios are low (0.25 to 0.05). Figure 12.8 displays some sample surfaces used as radiators. With the exception of the thinnest Teflon, all the surfaces emit between 0.6 and 0.80. The solar absorbances are low for all of these samples.

Figure 12.8: Sample Radiator Surfaces [12.12]

Heaters are commonly used to maintain an adequate temperature in all devices. The heaters can vary from less than 1 W in very sensitive devices to greater than 40 W. They are usually paired with a backup of equal power. Heaters are typically made of Kapton. Along with the heaters there are thermostats present to measure the temperature and turn on or off the circuit. A thermostat is designed with a set point and a dead band. The set point is the temperature or temperature range the heater is activated and the dead band is the range from the on and off temperature. All heaters and thermostats are wired in such a way that all thermostats must fail before the whole circuit will fail. The current industry standard for thermostats is solid-state controllers which use an electrical switch to turn on the heater circuits. They are more reliable and longer lasting than older mechanical thermostats. They are capable of using remote temperature sensors, small dead bands (< 4°) and function at freezing temperatures. Figure 12.9 shows a solid-state controller and its remote temperature sensor.


Final Report 140

Figure 12.9: Solid-state Controller [12.13]

Conductive isolation is a passive control method that is prevalent in items that are very sensitive to temperature changes such as mirrors and sensors. Isolation is also used to isolate propellant to decrease the heater power needed. Ideally, if hydrazine was perfectly isolated and no energy was lost, there would be no need for heaters. Titanium is used in this role because it has a low conductance of energy. If heat builds up on one end, the other end will not heat up as quickly as many other metals. 12.4 – Trade Studies and Decision Justification The first thermal challenge dealt with the combination of a completely enclosed piece, namely the core and the modularity of the modules. The modularity forced each piece to be thermally independent from everything else. Each module was responsible to reject its own heat. A monolithic satellite could employ a heat pipe that extends through all parts of the satellite and is linked to a large deployable radiator. It could collect all excess heat and transport it to the radiator. Trying to use a heat pipe on a modular system would be impractical because of the docking interactions. A completely enclosed piece is then left with a surface radiator. Cooling would be too expensive and bulky for the small thermal load that is presented by the core. If the core tried to radiate heat outward towards the modules and the other segments, the radiator would be able to absorb the energy coming from the modules. Therefore, there would be no net heat loss and maybe an increase in heat because some modules used more energy than the core. Thus, the core was extended 50 cm to provide enough surfaces to radiate the energy. The second challenge dealt with trying to minimize cost of the modules. Designing a universal module that could deal with any electrical load from 0 to 300 watts would be a lower cost than individually designing each module. Each module would have the same insulation, radiator surface, heaters and thermostats. This was attempted for the first few designs of the module thermal systems but every attempt was not successful. There is too much difference between the modules. The module with the reaction wheels used 200 W during the whole orbit while the Earth viewing telescope used 180 W while in the Sun. If a ratio of α/ε = 0.1/0.4 is used, the environmental load is 97 W in the Sun and 66 W in the dark. With these loads one module’s temperature will exceed the limits. Therefore, four separate module designs were created for the six different electrical load combinations that are present in the modules.


Final Report 141

12.5 – System Design The system design is broken up into three parts: the multilayered insulation, the propulsion segment and the optical segment

12.5.1 – Multilayered Insulation There was only one MLI designed for this mission. All other instances requiring insulation used low emittance tapes that are easily applied. Table 12.4 presents all the pieces that make up the MLI chosen for this design

Table 12.4: Insulation Layers

Name # of

LayersThickness mm (mil)

Weight (kg/mm2) α ε

Outer Layer

Kapton (Aluminum on backside)

1 0.254 (10) 0.036 0.44 0.62

Middle Layers

Aluminized Mylar 10 0.00635 (.25) 0.0093 0.12 0.03

Inner Layers


1 0.0254 (1) 0.05 NA 0.06 Aluminum side 0.4 reinforced side

Spacers Dacron 11 0.1651 (6.5) 0.0065 Total 12 2.159 (85) 0.2483 <0.01

12.5.2 – Propulsion Segment The propulsion segment is wrapped in Aluminized Kapton tape (α/ε = 0.31/0.45). [12.14] The surface is allowed to fluctuate from -23 to 77 °C because the surface is isolated from the internal components via titanium struts. The propellant tanks are surrounded by the MLI. The MLI has an emittance of 0.01 and emits 10.5 W. The six primary heaters on the tanks are made of Kapton and are rated at 5 W each. There are six secondary heaters and 12 solid-state controllers on each tank. If all of the heaters were to fail, s full tank would freeze in 30 days. The helium tanks are also wrapped with the MLI and emit 6.2 W. There are six primary heaters rated at 3 W each on the helium tanks and six secondary heaters along with 12 thermostats. There are 13 primary and 13 secondary heaters on the 12 thrusters and the arc jet. There are two thermostats for each heater. All of the thermostats have a set point of 10° C and a 2° C dead band. The propellant lines are covered with copper tape with an emittance of 0.02. There is one heater rated at 1 W every 2 m for the 90 m of propellant lines. There is a secondary heater in between each primary heater. There are two thermostats at each primary heater. The arc jet uses 1 kW of power and is dissipated through the propellant and the radiative emissions from the nozzle. [12.15]


Final Report 142

12.5.3 – Optical Segment The core has insulation around the surfaces that are covered by modules or segments. The exposed end is an aluminized Teflon radiator (α/ε = 0.14/0.4). [12.16] The modules are covered with insulation on the five surfaces facing the core or surrounding modules. The outer surface is a radiator. There are two heaters and two thermostats in each module. The power of each heater varies and is listed in Table 12.5. The super-module and the antenna module require the heaters to be on all of the time. The antenna has the lowest electrical consumption and mass of module parts. Therefore it is more sensitive to the changing environments from Sun to shadow. The space viewing telescope has a short baffle with five baffles over a 50 cm tube. The inside is painted with Chemglaze Z306 black paint (α/ε = 0.96/0.91). [12.17] The baffle is made of titanium with the MLI around the outside.

Table 12.5: Module Surfaces, Heaters and Temperature Range

Similar modules Item

Mass (kg)

Watts (Sun)

Watts (shadow)

Surfacenormal

vector (ρ)Surface material α ε Heaters

TemperatureRange (C)

1.2, 3.2 Computer 34 100 50 135 Aluminized

Teflon 0.14 0.4 15 W 15.4 18.1

1.3, 3.3 Antenna 12 5 5 135

Brilliant Aluminum

Paint 0.3 0.31

30 W (all the time) 10.7 19.2

2.2, 2.3 EVT 490 180 0 180 Aluminized

Teflon 0.14 0.4 15 W 11.9 12.6

1.1, 2.1, 3.1 RSU 12 40 40 135

Brilliant Aluminum

paint 0.3 0.31 30 W 10.5 13.9

1.5, 3.5 RWA 100 200 200 135 Silvered Quartz 0.070.79 15 W 12.5 13.2

1.4, 2.5, 3.4 Batteries 135 20 20 135

Brilliant Aluminum

paint 0.3 0.31 30 W 9.7 10.1

Super- Module SVT 800 70 70 0

Aluminized Teflon 0.14 0.4

30 W all the time) 7 10.2

The struts that support the CCD chip assembly on the SVT use a piezoelectric shape-memory material that, when heated, changes shape into a "parent shape" determined during manufacturing. This process is described in further detail in section 4.4.2.2 of this paper. When the SVT is ready to use, the struts must be heated to approximately 166 °C via wires running through them. This would cause the struts to return to its "parent shape". There are three struts that extend to meet at a focal point 6.875 m above the base of the segmented Beryllium mirror. The struts are each 9.65 mm in diameter with a 3 mm hole running through them. Therefore the wire that runs through these holes must have a diameter of less than 3mm.


Final Report 143

12.6 – Cost Estimation Table 12.6 lists the mass of all the components in the Thermal Control System. The mass of the insulation is calculated from Table 12.5. Heaters are made of DuPont Kapton XC which is listed as 1.41 gm/cm2

. [12.18] Using an average area of the heaters from Hubble, an area of 2.5x25 cm

was used. [12.19] The thermostats are Tayco solid-state controllers which are used on Hubble and the ISS. [12.13]

Table 12.6: Total Thermal Mass and Power

Device Weight

(kg) Power

(W) Back upQuantity

PrimaryQuantity

Surface Area (m2)

Total Mass(kg)

Total Energy

(W) Propulsion tank heaters 0.088 5 18 18 3.17 90 Propellant line heaters 0.088 1 45 45 7.93 45

Thruster heaters 0.088 1 13 13 1.15 13 Propellant thermostats 0.030 0.1 152 4.56 15.2

Propulsion tank insulation 0.248 3 2.92 2.17 0 Helium tank heaters 0.088 3 18 18 1.59 54

Helium tank thermostats 0.030 0.1 36 1.08 3.6 Propulsion segment outer surface 0.009 1 30.63 0.28 0

Module thermostats 0.030 0.1 30 0.9 3 Module heaters (large) 0.088 30 9 9 1.59 270 Module heaters (small) 0.088 15 6 6 1.06 90 Module inner insulation 0.248 14 3.79 13.19 0

Core insulation 0.248 1 38.03 9.44 0 Super module outer surface 0.020 1 80.2 1.60 0 Module outer surface paint 0.240 8 1 1.92 0

Module outer surface radiator 0.020 4 1 0.08 0 Quartz mirror 1 2 1 2 0

Final Weight (kg) 53.70 583.8 12.7 – Summary The preference of thermal system is to use passive components whenever possible. This system relies on many passive surfaces, insulation and isolation. The segment used insulation around the pipes, tanks and outer segment surface. There was plenty of titanium to minimize the thermal conductance between the parts of the propulsion segment. Heaters are used to maintain the temperature of the of the tanks. The modules were isolated from the other pieces around them because they used titanium connections and had five surfaces covered with 12 layers of insulation. The sixth surface has a radiator of Aluminized Teflon, aluminum paint or quartz mirrors. There are heaters and thermostats present to keep the systems functioning.


Final Report 144

I learned three things during this design course. First, I learned about radiative heat transfer and thermal control systems. I had no classes that dealt with it directly but I used a small amount of thermodynamics and spacecraft dynamics to help create my design. I learned enough to be able to talk with a professional thermal engineer from Boeing about more than just the basics of thermal control. The bottom line is I liked it a lot and was able to get excited about it. You cannot wait to design your system until you have your inputs from other groups. I wanted a lot of metrics about module sizes and power uses before I would design them. It was not until very late that I learned my environmental model needed higher fidelity and that I did have the ability to do it. I still think I could do more with it, with enough time. The last thing I learned was about the effects of unclear/impractical expectations, nit-picking and inconsistent instructions on the morale of a design team. As I write this, I am unsure if this meets the expectations of the course. The whole design team is unclear about this. That leads me to wonder if I am making a wise and full use of my time by typing a report that may or may not be what is needed. This possibility upsets me because I am proud of this report and because I have spent a great deal of time creating it. Nit-picking and inconsistent instructions frustrated this whole group because we were never right and there were few signs our efforts were moving towards the goal. Not all tried to improve, but those that did were frustrated with their lack of progress.


Final Report 145

12.8 – References [12.1] Brown, Charles D. Thermal Control. Elements of Spacecraft Design, 30 Dec. 2001,

AIAA. Castle Rock,CO: Wren Software, 200. [12.2] Brown, Charles D. Thermal Control. Elements of Spacecraft Design, 30 Dec. 2001,

AIAA. Castle Rock,CO: Wren Software, 200. [12.3] Gilmore, David G., ed. Spacecraft Thermal Control Handbook Volume 1: Fundamental

Technologies. 2nd ed. Vol. 1. El Segundo, California: The Aerospace Press, 2002. 23. [12.4] Gilmore, David G., ed. Spacecraft Thermal Control Handbook Volume 1: Fundamental






Technologies. 2nd ed. Vol. 1. El Segundo, California: The Aerospace Press, 2002. 148. [12.10] Brown, Charles D. Thermal Control. Elements of Spacecraft Design, 30 Dec. 2001,

AIAA. Castle Rock,CO: Wren Software, 200. [12.11] “Fine Guidance Sensor Mirror Cover." Swales Aerospace. 2 Dec. 2006

<http://www.swales.com/products/specialty_tool/fgs_mirror_cover.html>. [12.12] Gilmore, David G., ed. Spacecraft Thermal Control Handbook Volume 1: Fundamental

Technologies. 2nd ed. Vol. 1. El Segundo, California: The Aerospace Press, 2002. 791 [12.13] Gilmore, David G., ed. Spacecraft Thermal Control Handbook Volume 1: Fundamental


Technologies. 2nd ed. Vol. 1. El Segundo, California: The Aerospace Press, 2002. 793. [12.15] Hrbud, Ivana. Personal interview. 31 Oct. 2006.


Final Report 146

[12.16] Gilmore, David G., ed. Spacecraft Thermal Control Handbook Volume 1: Fundamental Technologies. 2nd ed. Vol. 1. El Segundo, California: The Aerospace Press, 2002. 791.

[12.17] Gilmore, David G., ed. Spacecraft Thermal Control Handbook Volume 1: Fundamental

Technologies. 2nd ed. Vol. 1. El Segundo, California: The Aerospace Press, 2002. 792. [12.18] "DuPont Kapton XC." DuPont. 2006. 2 Dec. 2006

<http://www2.dupont.com/Kapton/en_US/assets/downloads/pdf/XC_H-78314.pdf>. [12.19] Gilmore, David G., ed. Spacecraft Thermal Control Handbook Volume 1: Fundamental

Technologies. 2nd ed. Vol. 1. El Segundo, California: The Aerospace Press, 2002. 102-10


Final Report 147

13.0 – Experiment Segment The experiment segment, shown in Figure 13.1 is designed to house new modules containing technology demonstrators, university and industry experiments, and other yet to be conceived projects. By allowing for up to 16 modules to dock to the experiment segment it is possible to greatly expand the capabilities of Odyssey.

Figure 13.1: Experiment Segment

13.1 – Scope and Purpose The purpose of the experiment segment is to provide space, power, and data transfer capabilities for future expandability. This will allow for a highly modular and reconfigurable system with numerous mission capabilities. The type and limit of modules is defined only by the dimensions provided and imagination of the customer. Potential technology demonstrators and could include items such as power beaming technologies, directed energy weapons, docking stations for nanosats, and detections systems for ELE objects. Technology demonstrators would not be limited to these demonstrators, but these are some potential, and highly probably demonstrators. 13.2 – Power Beaming Demonstrator One potential technology demonstrator for Odyssey’s experiment module would be a power beaming demonstrator. A demonstrator could be developed that would allow for power to be beamed from the module to a nearby nanosat at regular intervals to demonstrating feasibility. Such a demonstrator would be extremely beneficial for future fractionated space systems and as a low cost feasibility study of this type of system.


Final Report 148

The power and data transfer capabilities of Odyssey allow for this type of system to be performed at low cost, with less expensive hardware. Both the nearby nanosat and module could communicate directly to ground controls and power data could be seen in near real-time allowing for problems to be detected and trouble shot without expensive antenna and processing capabilities.[13.1] 13.3 – Directed Energy Weapon A somewhat controversial system popular within the military arena is that of a directed energy weapon. A directed energy beam could potentially blind, or jam, or destroy a satellite eliminating a current threat, or as the precursor to a ground force. This would be beneficial as a space based system because there is no limit to the location of the satellite to be targeted. This type of system is being pursued by other countries, such as China, and will definitely pose a threat in the future. Although a full scale system would surpass the capabilities of Odyssey’s power and size capabilities a small scale demonstrator could easily be mounted within a module and stay within the required power limits. 13.4 – Nanosat Docking Station A large advantage of Odyssey and its modularity is the ability of the system to be a base station for a barrage of nanosats. Both hydrazine and helium lines traverse the length of Odyssey and have docking ports for additional segments. Because of this fuel transfer capability a segment could easily be designed to allow for the refueling of a Helium and/or hydrazine propelled nanosat. The current docking mechanism would allow for power to be transferred to a docked nanosat and data to be transferred wirelessly between the systems. The wireless capabilities of the current system would allow for nanosats operating within 30 km to communicate with and transfer data between odyssey, greatly reducing their need for large, power intensive computing and transmission systems. This overall concept allows for multiple expanded missions with quick response capability. Nanosats placed in orbit near Odyssey could quickly respond on request, then return to Odyssey to refuel and recharge while removing the need for large fuel tanks, power generation, and large computing systems. 13.5 – NEO and ELE object detection A potential application for a future module on the experiment section would be an ELE and NEO detection system. By attaching an automated radar or optical system within a module, space could be scanned for NEO’s. This type of system could help fill the current gaps in space based asteroid detection and allow for some warning, however brief. This type of system could be fully autonomous requiring no human intervention unless an object was found. The power and data requirements of this type of system should be within the capabilities of Odyssey; the ability


Final Report 149

for modular expansion makes it the ideal location for the start of a space based ELE object detection system. In 2006 there were 56 NEO approaches closer than 5 lunar diameters from Earth. Although none of these were ELE objects, any impact would still have a significant effect on the planet, and on day to day operations of many people. Being able to detect some of these and larger objects is and will be critical to the future.[13.2] 13.6 – Summary Odyssey’s experiment segment has room for 16 one meter long modules each of which could house a different type of experiment or technology demonstrator. Modules can be designed by entrepreneurs, industry partners, universities, the military, or anyone else with the means and desire for such a system. Each module will be low cost, make use of existing hardware on Odyssey and be able to be launched quicker than most systems today.


Final Report 150

13.7 – References [13.1] Turner, Andy. "In-Space Power Transfer Bouncing the Light - Fantastic!" DARPA.

Fractionated Spacecraft Workshop. Colorado Springs, CO. 4 Aug. 2006. 1 Nov. 2006. [13.2] "Near Earth Objects." Jet Propulsion Laboratory. 28 Nov. 2006. National Aeronautics and

Space Administration. 28 Nov. 2006 <http://neo.jpl.nasa.gov>.


Final Report 151

14.0 – Cost and Scheduling Determining the cost of a system is a large part of designing a system. Costs can determine how soon the system will be operational. Costs can also limit the technologies or materials used in a system. Finally, system costs can determine whether the construction and operation of that system is possible. Determining the timeline of a system is as important as cost when it comes to the success of a system. The timeline sets base goals or dates for milestones to be accomplished by. These milestones must be obtainable and allow for the completion/operation of the system in the desired timeframe and budget. For example, if mission operation is desired one day after the concept is realized it would be the part of scheduling to illustrate that such a timeline for construction, testing, and launch is not possible. 14.1 – Scope and Purpose The purpose of cost and scheduling is to determine a realistic cost estimate for The Odyssey Project and to establish a timeline for Odyssey’s development, construction, deployment, and service. The total system cost will be formed by summing the costs incurred during the timeline established by scheduling. The timeline that has been developed is feasible and allows for the mission parameters to be met. Costs were determined on a component and weight basis. Launch vehicle costs and launch operation costs were provided by SpaceX.[14.1] This cost analysis takes into account the costs of large items, such as power generation, propulsion, structures, SVT, EVT, and computing. A detailed list of parts and prices is beyond the scope or intent of this analysis. It is desirable that the cost and scheduling sections work together with all the other development sections to create a cost efficient timeline that allows for the customer’s goals to be met. 14.2 – Cost Analysis There are several major components that can drive the cost of a space system. The size of the space viewing mirror was a major driving factor in this design. The size of the mirror not only affects the weight of the system but also the size of the structure that must support it which in turn affects what launch vehicle is required. The fairing size that is required and the launch mass drive the selection of the launch vehicle. Launch vehicle costs are generally greater than fifty million dollars. In addition to the mirror size, there are other components that individually account for twenty to twenty-five percent of the vehicle cost. One of these components is the structure of the system which costs roughly twenty-five percent of the total system. Another component is the computer programming and software required to operate the system which incurs roughly twenty-one percent of the vehicle costs.


Final Report 152

The choice of the components and technologies used also greatly impact the cost of the system. The use of unproven or low TRL technologies is undesirable. The use of these technologies is undesirable because they are more likely to fail in orbit than proven technologies are. Failure in orbit means that these components will need to be replaced which causes costly additional launches. Furthermore, these low TRL technologies cost more to implement. For these reasons, components that are already in production are used wherever possible. In addition to vehicle construction, there are several other costs that appear. Once the vehicle is assembled, it must be tested. Testing has a time frame and price associated with it (the scheduling of which will be later discussed). Purchasing insurance for the system also leads to a substantial cost. Insurance costs are up from nine percent of the system’s cost in 1999 to sixteen percent in 2006. [14.2] The insurance cost for The Odyssey Project is almost seventy million dollars. Table 14.1 shows component costs for the major systems of the spacecraft. It also shows the launch and insurance costs.

Table 14.1: Component Costs Component Cost 1. Payload 1.1 Visible $43,969,601.60 1.2 Visible $72,396,832.18 1.3 Comm. 1.4 Antenna $1,366,900.56 1.5 Comm. electronics $14,678,000.00 2. Bus $105,306,853.73 2.1 Structure/thermal $5,146,367.86 2.2 TT&C $9,971,493.78 2.3 Attitude control 2.4 Attitude Deter. $18,263,955.30 2.5 Reaction Control $28,046,271.62 3. Power $1,919,843.71 4.Propulsion 4.1 3-axis stabilized $1,248,000.00 5. Aerospace Ground Equipment $31,405,000.00 6. Software program level $90,000,000.00 7.Testing (entire system) $12,000,000 8. Primary launch $51,000,000 9. Secondary launch $35,000,000 10. Secondary launch propulsion system to satellite $2,000,000 11. Insurance (16% of satellite cost) $75,098,763 12. Profit $62,708,233

Price tag to customer $653,533,138 Table 14.1 shows the total cost to the customer to launch the completed satellite without the experiment segment. This price includes the first two launches to get the satellite into orbit, functioning, and able to operate the space and Earth viewing telescopes. Component cost calculations were made using SMAD and can be found in the appendix.[14.3] An insurance cost of 16% of the satellite’s cost was given from the Futron Corporation. [14.2]


Final Report 153

The actual profit from a system is a very complex analysis because competitors always exist [14.4]. The profit generated could be greater or less than listed. The profit listed was calculated by adding an additional 10% of the total system cost. This percent can be raised or lowered to make the bid more competitive. For the system to completely meet the operational requirements a third launch will be required. This launch will include the experiment segment that will attach to the orbiting space system. This experiment segment can be launched filled with modules or empty which requires that individual modules be launched separately later at a greater launch cost to the customer. Tables 14.2 and 14.3 show the costs for customers associated with these two launch conditions.

Table 14.2: Costs to launch multiple modules Third launch if 16 modules used Launch vehicle (Flacon 9) $27,000,000Propulsion to satellite $2,000,000Cost to orbit $29,007,023Added 16% insurance $33,648,146Profit $3,364,815Total cost to customer $37,012,961

Table 14.3: Costs to launch single module

Third launch for single modules Launch vehicle (Falcon 5) $18,000,000Propulsion to satellite $2,000,000Cost to orbit $20,006,053Added 16% insurance $23,207,021Profit $2,320,702Price tag to customer $25,527,723

A comparison between Table 14.3 and Table 14.4 shows the saving to the customers who do not require an immediate launch. Table 14.3 shows that launching a single module costs $25.5 million. Launching a full experiment segment costs $2.3 million per module. Launching an experiment segment with 16 modules results in a cost savings of $23.2 million per module. The customer will pay a flat cost of $690.5 million for the fully operational, completed system. The customer should expect to pay a re-fueling cost of $5 million every year for the operational life of the satellite. The customer is also responsible for the operational costs associated with the spacecraft, such as ground facilities, data transfer, and third user relations. Recommendations on these areas are further discussed in the concept of operation.

14.2.1 Hubble Comparison Hubble cost $1.5 billion to build and put into orbit. The most recent of five repair missions is estimated to cost $900 million. The Odyssey Project has a projected construction and orbit insertion cost of $690.5 million. It will also be possible to launch individual replacement modules for $25 million each. The modularity of The Project Odyssey enables individual systems to be replaced or upgraded with a single launch. Furthermore, repairs or replacements


Final Report 154

do not require a shuttle launch which can be very costly. Hubble has been operational for 16 years and is planned for operation through 2013. The technology is available and proven for a spacecraft to last over 20 years. 14.3 – Scheduling A schedule that is created for the project must be feasible. The schedule can not create a time frame for the completion of tasks that is impossible. Restrictions such as material availability, construction facilities, assembly time, and testing time must all be taken into account in order to make a reasonable schedule. Figure 14.1 shows the proposed timeline for The Odyssey Project. This timeline is consistent with design and operational timelines of monolithic satellites. However the construction and testing of The Odyssey Project is shorter than industry standards by four years. This shortened time for construction and testing is made possible by the modular design of the system. Each sub-system (power generation, propulsion, guidance) can be manufactured and tested independently of the other sub-systems. A final assembly and test is still required however this test can be far less intensive due to the sub-systems’ previous testing. This shortened construction and testing has enabled The Odyssey Project to meet the customers request for a primary launch by 2011.

Figure 14.1: Overview of project timeline

The Odyssey Project begins with the request for proposal in the first quarter of 2006 calendar year. The design is finalized in the third quarter of 2007. Next, construction begins in the fourth quarter of 2007 and continues through the first quarter of 2009 when testing begins. Testing ends in the fourth quarter of 2010. The primary launch which does not include the super module or the experimental section is scheduled for the third quarter of 2011. This primary launch includes the propulsion segment and the optical segment without the super module. The second launch, scheduled for the fourth quarter of 2011, will deliver the super module. Finally, additional launches are scheduled to begin in 2015. These additional launches will deliver the experiment segment with technology demonstrators to the satellite. The EVT will be operational in 2011 with the SVT following in 2012. Requests for time on either system can be made in 2010 and 2011 respectively. Ideally for the customer will begin taking requests for technology demonstrators by 2013. This will allow two years for third


Final Report 155

parties to develop their technology demonstrators before the scheduled launch of the experiment segment in 2015.


Final Report 156

14.4 – References [14.1] “Revolutionizing Access to Space.” SpaceX. 15 Nov. 2006. < http://www.spacex.com/> [14.2] Futron Corporation. 2002. “Satellite insurance rates on the rise - market correction or overreaction?” [14.3] Wiley J. Larson and James R. Wertz. 1993. Space Mission Analysis and Design, 3rd Edition. Microcosm Inc. [14.4] Seokcheon Lee. Personal Interview. 20 Nov. 2006.

http://www.spacex.com/


Final Report 157

15.0 – ConOps The concept of operations (ConOps) is a cooperative venture between systems engineering and cost and scheduling. ConOps details the operational concept and overall costs to the company highlighting the benefits and experiences. In addition to creating a viable operational plan, ConOps also performs mission planning and operations. 15.1 – Scope and Purpose The purpose of ConOps is to develop a comprehensive operational plan for the company to allow for a profitable business and to develop the system operational and future plans. This section will discuss ground facilities, operational plan, and future expandability of the system. 15.2 – Business Plan The total cost for Odyssey will be $652.9 million for design, material cost, construction, testing and launch. There are three primary methods in which the customer can use Odyssey to generate revenue, SVM, EVC, and technology demonstrators. The customer can sell use of its SVM to third parties to generate Hubble like images of space. The use of the EVC can also be sold to third parties for the competitive Earth imaging market. The income generated by selling the use of the SVM and EVC should generate sufficient revenue. Technology demonstrators can also be sold by the customer to third parties in one meter modules for launch on a Falcon 5 vehicle, or at a lower cost if 12-16 modules are launched as a segment onboard a Falcon 9. The third party could then be charged a per month rent for the space their technology demonstrator occupies and the power that it requires to operate. Additional single module launches will cost the customer $25.5 million, or $37 million for a segment launch. 15.3 – Ground Systems The ground systems for The Odyssey Project will be based in Purdue Research Park in West Lafayette, Indiana as shown in Figure 15.1. The facility will include offices for 25 employees, two conference rooms, and research space for an additional two employees. All communication with Odyssey will be done via NASA TDRSS sent from the West Lafayette facility or an alternate site at the customers choosing. This will allow for a single facility operation keeping project overhead and facility costs to a minimum. Communication with NASA TDRSS will be done via the internet and secure servers to the located at the NASA communications center in White Sands, New Mexico.


Final Report 158

Satellite construction will take place in or near Vandenberg Air Force base in California. All operational launches will be from the SpaceX launch facility at Vandenberg Air Force base. As such a facility near the launch site is a possible future expansion for better coordination. The customer should plan on maintaining all operations with Purdue Research Park with constant communication links between NASA TDRSS and the SpaceX launch site at Vandenberg Air Force Base or by constructing their own facility at an alternate location.

Figure 15.1: Purdue University Research Park[15.1]


Final Report 159

15.4 – Operations The operational component of the customer will be determining the timetable and scheduling with the third party to increase profits while providing quality benefits suitable to each customer. The operational plan can be broken down into:

• What the customer provides • What the third party wants • Contract with the third party • Date and Time scheduled • Task performed • Information provided to third party

These six steps outline the basics for each of the three profit generating components of The Odyssey Project.

15.4.1 – Space Viewing Telescope The operational process for the space viewing telescope will begin with the third party requesting time on the telescope to view a certain object in space. The third party will provide the exact location of the object to be viewed as well as the desired time and date. The customer will work with the third party to set a date and time on the schedule that will allow for proper viewing of the object without interfering with other satellite operations. Once the time and date has been set it will be verified with the satellite operations group (SOG) to make sure that the satellite can be maneuvered into the proper orientation to view the object. The SOG will then determine the maneuvering sequence to properly orient the satellite and place this in the system to be processed at the correct date and time. As the time to view the object arrives the SOG will recheck the queue and maneuvering commands and then pass them to Odyssey through TDRSS. Odyssey will perform the proper maneuvers and take the requested images transmitting the raw data back to the operations center via TDRSS. After the requested images have been taken, Odyssey will maneuver into position for the next set of images. After the data has been received on the ground, the data will be compiled and sent in either compressed or raw format depending on third party request. The third party will then receive the data and make any additional image requests of that object before their satellite time has expired.

15.4.2 – Earth Viewing Telescope For the Earth viewing telescope the customer will provide the place to be imaged, and only provide specific times on rare occasions such as natural disaster, or time sensitive military operation. Once the location has been determined and verified, the customer will work with the third party to schedule time on the telescope.


Final Report 160

The customer will then work with the SOG to get the Earth viewing telescope in the proper position and instruct the telescope via TDRSS to take the requested images. After the images have been taken they will be delivered in raw data format to the ground station via TDRSS. The images will then be processed and sent to the third party.

15.4.3 – Technology Demonstrator For the third party to send a technology demonstrator to Odyssey they would first express their interest to the customer. The customer would then provide the third party with a detailed design rules packet including module dimensions, power limits and data transfer requirements. The third party would then design the system and work with the customer to finalize the design and order all required hardware. All hardware would be shipped to the launch processing facility where contractors would perform the assembly of the module under the direction of both the customer and the third party. Once assembly has been completed a detailed testing schedule will be performed to verify connections and communications capability with Odyssey. The module will then be attached to the MPS and packaged into the payload fairing with other modules. After full fairing assembly the payload will be connected to the launch vehicle and moved to the launch pad for launch. After a successful launch and orbit insertion the payload fairing will separate and the modules with MPS will be free floating in space. Odyssey will then direct the modules into proper docking order with the satellite. After docking has been achieved and connections verified the third party will instruct the customer to activate the module. After the module has been activated all data will be transferred to the ground station where it will be relayed to the third party. Some data processing may be available depending on the nature of the data and experiment or demonstration. 15.5 – Difference between fractionated systems and single function system Operationally there is a significant difference between a single monolithic system and a fractionated system. A single monolithic system requires a larger amount of redundancy but will remain virtually unchanged throughout its lifespan. This differs significantly from a fractionated system where the satellite is constantly undergoing changes in size, shape, and function. A single monolithic system would require fewer ground station personnel due to the unchanging status of the system. There would be no need for additional launches, or upgrades and all systems would remain identical to their original form. This means there are fewer operational functions as an entire section of adding modules would be gone. A fractionated system is more complex and requires more operational functions. By adding the complexity of attaching and detaching components and modules the number of people to insure


Final Report 161

successful operation also increases. Although the cost does increase significantly with increased personal and launches, the additional revenue stream can offset this. 15.6 – Summary Overall the system will be functional in support of the aforementioned operational plan. The modularization of the system will create an additional revenue stream allowing for higher performance system with greatly expanded capability. Working with each third party to ensure that they receive the information that they desire will allow for the customer to prosper and for all three systems, space viewing, earth viewing, and technology demonstrators to generate revenue.


Final Report 162

15.7 – References [15.1] Purdue Research Park Site Map <http://www.purdueresearchpark.com>


Final Report 163

16.0 – Conclusion The Odyssey Project is a reconfigurable, upgradeable, and modular system capable of Earth and space imaging comparable to the Google Maps and the Hubble Space Telescope. The system is a blend of current tried and tested technologies with a high TRL, and newer advanced technologies. 16.1 – Project Summary Odyssey will orbit the Earth in a polar orbit of 98.5° at an altitude of 785 km retracing each path every 25 days. The maximum space between ground tracks will be 113km within the capabilities of the pivoting Earth imaging system operating with a 1m resolution. The altitude for desired orbit and the orientation of Odyssey will be maintained via an electric propulsion engine and decomposition thrusters. The propulsion system uses hydrazine and helium as fuel and pressurant providing for refueling capability via Orbital Express. The thrusters act in conjunction with the reaction wheels, rate sensing units, and fine guidance sensors to achieve the proper orientation of Odyssey and to ensure a 0.05 arcsec accuracy for the SVT. The SVT has a resolution equal to that of Hubble but has a greatly expanded computer processor and data system. This radiation hardened 2.2 GHz system is capable of transmitting data to Earth via TDRSS on Ku and S bands using a dual antenna system. In conjunction with these optical capabilities the system has an important modularity function. Most components on the system can be replaced via the MPS as individual or groups of modules. Additional modules can be added to the experiment segment using the MPS increasing the functionality of the system. Each MPS is controlled by the ACS onboard Odyssey after launch and is docked using a LADAR system installed on the core. Each system on Odyssey has the required thermal insulation and protection to maintain desired temperatures for operation. Several systems including the propulsion and power system have heaters installed to elevate the temperature to an acceptable level. The power system consists of 65 m2 of solar panels with 93 kg of NiH batteries for storage. The entire system can be launched from Vandenberg Air Force Base in three separate launches of SpaceX Falcon class launch vehicles. Total launch cost for all three launches will be $80 million, raising the total system cost to $653 million. This cost includes work from conception in 2006 through initial launch in 2011 and operational budget and future launches through 2021.


Final Report 164

Figure 16.1: Odyssey Orbiting Earth[16.1]

16.2 – Lessons Learned Throughout The Odyssey Project design process many things were learned that would have contributed to a different design than currently shown. As the suggested modular and fractionated system was foreign to most of the design team, there was little first hand knowledge and throughout most of the design process the required information was difficult to locate. When found, it was often too late for new information to be incorporated into the design. As a result of the lack of knowledge, some critical design decisions had to be made early on in the design process with less than desired information. As the design process progressed it became clear that some previous decisions were poor and had negative impacts on other systems. Perhaps a better method would be to choose a concept very early in the semester, leaving time for a redesign once more knowledge has been learned about the specific systems (Figure 16.2). This type of redesign allows for the design team to better learn their system sooner, and allows for poor initial decisions to be corrected within a new design. This would also allow for more overall progress with more time spent designing the system with knowledge in hand, and less time spent creating concepts without knowledge of systems and systems operation.


Final Report 165

Figure 16.2: Progress vs. Time plot

One of the most important lessons learned throughout the design process was the incorporation of outside sources, including experts in the field. Many individuals with relevant experience exist at Purdue and are willing to help when asked. Their detailed first hand knowledge of individual subsystems can be extremely helpful and lead to a better understanding of the materials. Throughout the semester it became obvious that not only did each team member have much to contribute, but was quickly becoming an expert in his or her area. Several team members were well suited for their group within the team, while others’ talents could have been better utilized on other teams. Overall, much was learned about each individual and if done again, the teams could be better organized to make use of the talent that already exists. 16.3 – Closing and Remarks In all, the project was successful in broadening the horizons of the students and introducing them to fractionated and modular space systems. The project forced everyone to throw out many preconceived notions and to “think outside the box” using new ideas and technologies. Each team member is leaving this class with a better understanding of space systems and an appreciation for the complexity of satellite systems.


Final Report 166

16.4 – References [16.1] "Earth 3D Space Tour." Software Laboratory. 2004. FP Software Lab. 11 Oct. 2006

<http://www.fpsoftlab.com>.


Final Report 167

Appendix


Final Report 168

A.1 - Team Biographies Hadi Ali Hadi Ali, a student member in AIAA, ASME, IEEE and SAE, is from Jordan and holds a B.Sc. in Mechanical Engineering (ME) from the University of Jordan. During his study, he worked at The Chair of Engineering Design at the University of Erlangen-Nürnberg, Germany on computer-aided design projects. His ME senior design project (a robotic cart) won a prize of appreciation from the Higher Council for Sciences and Technology in Jordan. After graduation, he worked as a Junior Engineer in a Liquified Natural Gas plant in Qatar. He joined Purdue in Fall 2005 as an Aero/Astro undergraduate student and will graduate in Fall 2006 with a major in Propulsion and minor in Design. In The Odyssey Project, he worked on Structures, Mechanisms and Materials. After graduation, he will purse graduate studies at Purdue specializing in Aerospace Systems Design. Hadi looks forward to contribute to the journey to Moon, Mars and beyond. Nick Andrews Nicholas Andrews will graduate with a B.S. in Aeronautical and Astronautical Engineering in December 2006. His main focus on The Odyssey Project was the Earth Viewing Telescope; this meant a detailed understanding of how a spacecraft orbits Earth, how it needs to view Earth and how to capture an image of Earth. He also worked with CATIA to visualize many of the parts designed for the payload team. His previous work experience has been with Northrop Grumman Defensive Systems Division as a systems engineering intern for the past 2 years. He has also worked with several on-campus organizations during his stay at Purdue. These include Purdue F.R.I.S.T. Programs, an organization that works with high school students to build and compete robots in regional and national competitions; he has also been a part of E.P.I.C.S. (Engineering Projects in Community Service) who builds small scale projects for charitable organizations. He is currently looking for a job in the Aerospace field, preferably something in spacecraft design or propulsion. Craig Bittner Craig Bittner is from Evansville, Indiana and came to Purdue in the fall of 2002 to Aerospace Engineering. He will graduate in December 2006. Within Aerospace engineering, his major area of focus is structures and his minor is aerodynamics. For this project, Craig designed the data handling system which involved find a modern processor capable of space application. Outside of class, Craig has been involved with Air Force ROTC and will commission as a second lieutenant in December after graduation. He will be stationed at Edwards AFB, California and work as a project engineer. Matt Dennis Matt Dennis is from Plano, Texas just north of Dallas. He came to Purdue University in the fall of 2006 to pursue a Bachelors of Science in Aeronautical and Astronautical Engineering. His major area of focus in aeronautics and astronautics was propulsion with a minor in dynamics. In Matt's areas of interest, he took classes on both the aero and astro disciplines. For The Odyssey Project, he was responsible for the design of the communication system and also contributed on


Final Report 169

several other systems. After graduation on 17 December 2006 he will return to Dallas and pursue a career in the aerospace industry. Jon Fromm Jonathan “Jon” Fromm is from Gurnee Illinois. He entered Purdue in the fall of 2002 and will be Graduating in December 2006. Jon is propulsion and design specialist who has worked as a test engineer at the Purdue Zucrow Laboratories. There he specializes in the design of Hydrogen-Oxygen propellant systems. Jon is exploring job options with several companies for testing positions related to manned spaceflight systems. Elisabeth Hanssens Elisabeth lived in California for the first ten years of life and then moved to Indiana. After graduation of high school in June 2003, she enrolled at Purdue University in August 2003. Following the first semester of Freshman Engineering, Elisabeth entered the School of Civil Engineering. During the fall of 2004, she worked as a co-op engineer in southern California for a construction (design-build) company. She changed her degree to Aeronautical/Astronautical Engineering when she returned from California. Within the school her major was propulsion and minor was design. She spent the summer of 2006 in Utah working for ATK as an intern. She worked on the ballistics and grain design team for the solid rocket boosters on the Shuttle. Elisabeth graduate in May 2007. Matt Harvey Matt Harvey is from Newburgh Indiana, he came to Purdue in 2002 after graduating from Reitz Memorial High School. His major area was aerodynamics with design as his minor. Matt will graduate from Purdue in December of 2006. He spent the previous three summers working an engineering internship at Project Associates Inc. in Evansville Indiana. David Helderman David Helderman is from Greenwood, Indiana and entered Purdue University in the fall of 2002 to pursue a degree in Aeronautical and Astronautical Engineering. While at Purdue he has concentrated his efforts on a major in spacecraft design and a minor in propulsion. He has worked at the Purdue University High Pressure Lab since May of 2006 performing the design, building, and testing of a facility to support methane-oxygen ignitor testing. David has also been a part of a separate design, build, and test project for a hydrogen-oxygen ignitor at the lab. On The Odyssey Project David was the co-team leader and a member of the propulsion team. After graduation in December 2006 he plans to stay at Purdue to pursue a masters degree in aerospace propulsion. Norman Herbertz Norman Herbertz is a 5th year senior at Purdue from San Antonio Texas with plans to graduate in December of 2006. His preferred area of study is in propulsion, but his contribution to the team lay within the power systems. This also an area of interest to him as his father is a master electrician and he has spent most of his life around electrical power systems. The past four summers have provided him with work at his father’s cellular construction firm. While this is


Final Report 170

not an aeronautical firm it has taught him value of hard work and commitment. His future plans will have him working in Texas if he finds a desirable employment there. April Miller April Miller is from St. Louis MO and started attending Purdue in fall of 2002 with plans to graduate in December of 2006. Her focus throughout the project has been in the area of structures operating as the structures team leader and primary developer of the docking process. April’s major area in AAE is structures and minor area propulsion, she has also worked toward an English minor. She plans to eventually work in the south in a position focusing on the structure of aerospace vehicles and to move steadily up in the company. Chris Murphy Chris Murphy is from Carmel, Indiana and enrolled at Purdue in the Fall of 2002 focusing on dynamics and propulsion for his major and minor. He will receive a minors in Political Science and Aerospace Studies. On The Odyssey Project he was in charge of designing the Thermal Control System. After graduation in December 2006 he will be commissioning into the Air Force as an Air Battle Manager with plans to make the Air Force a career. Ashley Ruic Ashley Ruic came to Purdue from a western suburb of Cleveland, OH in fall of 2002. Inspired by her experiences with the Explorer program at NASA Glenn Research Facility, she came to Purdue in hopes of performing undergraduate research. She was granted this experience when her team's research entitled, "The Effects of Sessile Rotating Magnerheological Fluid in a Changing Magnetic Field" was chosen to fly on NASA's micro-gravity testing platform, the "Vomit Comet." While this may have been the highlight of Ashley's student career, the Design major and Propulsion minor gained other valuable experiences in her time at Purdue. Through two summer internship sessions with Rolls-Royce, Indianapolis, she gained invaluable leadership experience in the aerospace industry. In addition, Ashley's time spent in Navy ROTC allowed her to apply her skills aboard the aircraft carrier, USS Stennis, with an F/A-18 squadron as well as with a CH-60 helicopter squadron in California. A December 2006 graduate of Purdue's AAE program, Ashley will be report to the Navy's flight school in Pensecola, FL in February 2007. Ashley's future goals include applying to the Navy's test pilot school, pursuing a masters degree in AAE, and possibly applying for the astronaut program. Aaron Schinder Aaron Schinder is an astronautical engineering major, with a major concentration in propulsion, a minor concentration in structures, and an economics minor. He will be graduating on 17 December 2006 and has in the summer of 2006 had an internship working with Dr. Ruggles-Wrenn of AFIT on a model for viscoelastic deformation. His area of technical design responsibility on the Odyssey project was the primary structural model for the core and modules. His future career will be in the Air Force with his first assignment at Kirtland AFB, NM as an aeronautical engineer.


Final Report 171

Nick Sochinski Nicholas Sochinski is a fifth year senior at Purdue University in the last semester of his Astronautical Engineering degree. At a young age Nicholas was always fascinated with flight and the possibility of being a pilot or perhaps an astronaut. As a student Nicholas soon found he had a passion for math and physics leading him into the obvious field of engineering. Although flight and aircrafts were very appealing to Nick, his thoughts dwelt on space and the opportunities it held. During the course of his degree, Nick found propulsion, alternative power sources, communications, and cost driven design being the biggest hurdles in the aerospace industry, thus he focused his attention on those areas. Along with Nick’s studies in engineering his is also a Cadet in Air Force ROTC from which he has held many leadership positions and been educated in leadership and group dynamics. As a senior Nicholas received a pilot slot from the Air Force, and after his graduation he will be sent to pilot training. Stephanie White Stephanie White is a senior in aeronautical and astronautical engineering at Purdue University. She entered Purdue in 2002 and plans to graduate in December 2006. She chose to concentrate on spacecrafts over airplanes because of her long fascination with the space program. Through her coursework, she has focused mainly on dynamics and controls. On the senior design project, her primary responsibilities were orbit selection and design of the attitude control system.


Final Report 172

A.4 – Payload – Optics A.4.1 – Earth Viewing Telescope (EVT) A.4.1.1 – Main Mirror Sizing The main mirror is the primary collection source for any telescope. It collects electromagnetic energy. This energy can then be reflected off the mirror and onto sensors that can detect it. The sizing of the mirror depends on several factors the most important two of which are the distance between the collector, mirror, and the energy source and the wavelength or frequency of the energy. Equation A.4.1 relates the distance from the energy source h, the wavelength λ, D is the diameter of the collector needed and X is the ground resolution.

XhD λ44.2

= ( 4.3) [A.4.1]

Equation A.4.2 is the angular resolution of a mirror; this allows us to calculate the greatest amount of area that can be seen with a telescope with a primary collector of size D. Where θr is the angular resolution in radians, D is the diameter of the collector and wavelength is λ.

Drλθ 22.1= ( 4.4)[A.4.1]

The electromagnetic spectrum as can be seen in Figure A.4.1, encompasses a huge range of frequencies. For the scope of the EVT, the project is only concerned with the visual and infrared range on the spectrum. The visual range starts at 350 nm and does to 800 nm. The infrared spectrum starts at 800 nm and ends at about 1000 μm.

Figure A.4.4: Electromagnetic Spectrum[A.4.1]

This information allows for a general idea of what the size of the mirror must be. There is one more piece of information that is needed. According to the RFP the Odyssey must take be able to take an image of any point on earth during its 25 day orbit. [A.4.2] This sets the wavelength of light


Final Report 173

to be looked at. The orbit, see Section 5, shows that the orbit only spends 30% of the time in the umbra of the Earth. This means that putting on an inferred system is not necessary; pictures can be taken in the 70% of the time the craft is in the sun. Also the orbit takes the satellite over the same point on Earth multiple times. If the target image is unable to be taken at the light conditions present on the first pass. Then it can be take later. This means that the spectrum of light to be viewed is the visual spectrum. The main region used in Earth observation satellites is the blue-green part of the visual spectrum which from Figure A.4.1 we can see at 500 nm. Using Equation A.4.1 with a λ of 5 nm, a resolution of 1 m and an orbital height of 800 km, the mirror size we need is 0.976 m. A.4.1.2 – Mirror Assembly The main mirror is only one part of the total telescope. To have a complete telescope system a set of lens or mirrors must capture the image from the primary collection surface and pass it onto sensors. This is done because having a sensor be the main collection surface is not feasible for most applications.[A.4.1] This allows for smaller sensors to do the same job as a one that is larger. The typical design of an Earth observation telescope is shown in Figure A.4.3. This design supports a primary parabolic primary mirror and a convex secondary mirror. This creates a compact telescope with only one major flaw; a large hole must be cut in the primary mirror to allow for light to pass through.

Figure A.4.5: Light Path of a Shmidt-Cassagarin Telescope [A.4.3]

This design was chosen over say a Gregorian type telescope Figure A.4.4, which has a parabolic primary mirror and a hyperbolic secondary mirror but suffers from a long focal length. Or a Nasmyth telescope which has no hole in the primary mirror allowing it to collect all light of a target area. But suffers from a large mechanical supports. A Shmidt type telescope is the design used currently for Earth observation type satellites such as IKONOS and Quikbird.


Final Report 174

Figure A.4.6: Gregorian Telescope [A.4.4] Figure A.4.7: Nasmyth Telscope [A.4.4]

The focal point for the primary mirror is calculated by Equation A.4.3. Where f is the focal length d is the detector length of 15 μm (discussed more in section 4.1.7), h is the height of the spacecraft and X is the ground resolution, as discussed in the RFP as 1 meter.[A.4.2]

Xhdf = (A.4.5) [A.4.1]

This gives a focal length of 0.6 m. To find the curvature of a mirror Equation A.4.4 must be used were D is the diameter of the mirror, f is the focal length and C is the curvature. The curvature required that that focal length to be 4.9m, easily within tolerances for the mirror manufactures.[A.4.1] Equation A.4.3 must be used again to find the curvature of the secondary mirror. The CCD chipset is placed 0.5 meters back from the main mirror. This is done because it is the farthest back that the geometry of the module allows. This is done to create a sharp clear image on the CCD chips. Using a diameter D of 0.205 m, a focal length of 1.1 m, the required curvature is 0.75 m. The results of this are seen in Table A.4.1.

fCD111

=+ (A.4.6)[A.4.4]

Table A.4.2: EVT Mirror Details Diameter (m) Focal Length (m) Curvature (m)

Primary Mirror 0.976 0.6 4.9Secondary 0.205 1.1 0.75

These mirrors must be held in place by a rigid support structure or the images produced by the telescope will be fuzzy and out of focus. Section A.4.1.6 discusses this structural necessity. A.4.1.3 – Mirror Materials The material the mirror is composed of is a very important design decision. Table A.4.2 below shows some common mirror materials used in orbital observation platforms.[A.4.5]

http://upload.wikimedia.org/wikipedia/en/6/64/Gregorian_Telescope_Lightpath.svg

http://upload.wikimedia.org/wikipedia/en/6/64/Gregorian_Telescope_Lightpath.svg


Final Report 175

Table A.4.3: Mirror Material Properties [A.4.5]

Brittleness density CTE(300) heat cond heat capy diffusivityMpa-m1/2 g/cm3 ppm/K w/mK j/kg-K 1e-6 m2/s

ULE 68 2.20 <0.03 1.31 776 0.8Zerodur 91 2.53 <0.05 1.46 821 0.8silica 73 2.20 0.52 1.38 703 0.84SiC 466 3.05 2.37 300 660 146Borofloat 63 2.22 3.2 1.1 830 0.6Pyrex 64 2.23 3.2 1.3 726 0.7

The material selected for use for the EVT on the odyssey was ULE (Ultra Low Expansion) glass. It was chosen based on mainly on Coefficient of Thermal Expansion (CTE). This is the main factor in designing mirror components.[A.4.5] because when glass expands radius of curvature decreases, this can be inferred from Equation A.4.4. Unfortunately ULE is a very brittle material, as Table A.4.2 shows. This means that either it must be made exceptionally thicker then required or have a support system to withstand the stresses imparted on the structure, further information on this will in section A.4.1.4. A.4.1.4 – Mirror Construction The primary and secondary mirrors of the Odyssey will be composed of ULE glass, and therefore it needs a support structure to withstand launch forces. This is also done to reduce the weight of the mirror. Figure A.4.5 shows how the EVT will accomplish this. The mirror will incorporate a series of aluminum struts to minimize deflection of the mirror. Figure A.4.6 shows how the mirror will deform at launch. This is the time when the maximum stress is applied to the mirror system.

Figure A.4.8: 3D Image of Honeycombed Mirror[A.4.5]

Figure A.4.9: Deformation with Max Loads (Launch)[A.4.5]


Final Report 176

Figure A.4.7 shows how the web plates of the honeycombed section of the primary mirror are arranged. The yellow webs are 5mm thick, while the red webs are thicken to 10mm. This allows the mirror to survive the required g loadings and still be able to complete its mission once in orbit. Section A.4.1.6 will detail the mass of the entire system including information on how these supports were computed.

Figure A.4.10: Aluminum Web plate Design[A.4.5]

A.4.1.5 – Stray Light Shroud Design The shroud of the EVT is divided into two sections, the outer surface of the shroud and the inner surface. The outer surface is designed to reflect heat and ambient light, while the inner surface is designed to absorb light and keep reflections from hitting the primary mirror. The interior of the shroud is composed of a honeycombed truss of aluminum beams to hold the shape of the shroud and to attach the skins of the inner and outer shrouds, and to hold the sun-shade in place. A.4.1.5.1 – Outer Shroud This section of the shroud has one function, to reflect heat and light away from the EVT. This is done for several reasons, first is to keep the mirror cool. If the mirror expands then it becomes unable to reflect light the way it should, causing the mirror to be distorted. To reflect as much light as possible the outer shroud is painted white. This reflects the maximum amount of energy away from the telescope. The outer shroud’s skin is composed of aluminized Teflon; this reflects large amounts of thermal energy away from the telescope. Aluminized Teflon is used as a primary thermal radiator on commercial applications; see section 12.5 for more information on thermal properties of this material and its use on the Odyssey.


Final Report 177

Figure A.4.11: EVT Light Reducing Shroud

The outer shroud also has a sunshade. This sun shade is used to cover any outside light sources other then the Earth. Outside sources include the Sun, the Moon or other spacecraft. This keeps the stray light from the mirror assembly.

Figure A.4.12: Sun-Shade

A.4.1.5.1 – Inner Shroud The purpose of the inner shroud is to keep stray light from reflecting onto the main mirror the inner skin of the shroud is made of aluminized Teflon which has had a flat back paint applied to the surface. This absorbs any addition light that is not part of the main image of the target location. In addition to black paint the interior of shroud is lined with baffles this reduces the stray light that enters into the main mirror.

Figure A.4.13: Baffles


Final Report 178

4.1.6 – Structural Support

Figure A.4.14: Mirror Assembly

The structure of the EVT was computed using the following equations. For simplicity the material was assumed to be solid 7075-T6 Aluminum with following properties.

Simple Rods FLmyield

*⎟⎟⎠

⎞⎜⎜⎝

⎛=

σρ (A.4.7) [A.4.6]

Beam Buckling 2

4

2

*L

chkEL

kEIFcr == (A.4.8)[A.4.6]

Beam Bending 4*12/12/*

hhM

IMy

yield ==σ (A.4.9)[A.4.6]

Table A.4.4: Al 7075-T6 Properties[A.4.5]

Ult Tensilie Strength 572 MPaTensile Yield Strength 503 MPaModules of Elasticity 71.7 GPa

Poisson's Ratio 0.33Shear Modulus 26.9 GPaShear Strength 331 MPa

Since the mirror is composed of ULE glass and Aluminum support structures the mass of the mirror was calculated from Equation to be 21.392 kg. Table A.4.4 shows the breakdown of the


Final Report 179

masses of the components in the EVT. Section 9 has more information on the structure of the Odyssey and how it was designed. All other masses were based on information gathered on existing systems and converted for use with the Odyssey spacecraft.

)0.38-)(D4

(D/8)(.875)-(1 22ULE

πρ=Mass (A.4. 10)

Table A.4.5: Mass Component Breakdown Odessey Earth Viewing Assembly

size or area Quantity Mass (each)

CBE Total (kg)

Optics (ULE & t = D / 8) % LtwtPrimary Mirror (0.38φ h 88% 0.98 1 21.392 21.392

Bipod Mounts to Optics Bench 3 9.500 28.500Secondary Mirror 80% 0.20 1 0.374 0.374

Tripod Focus Mechansim 1 11.000 11.000CCD 80% 0.32 1 1.435 1.435

Tripod Focus Mechansim 1 11.000 11.000Subtotal 73.70

Telescope StructureSecondary Mirror Housing & Struts 1 31.700 31.700Optics Bench Flat & Curbs 1 92.000 92.000Primary "Cold Stovepipe" Light Baffle 1 10.590 10.590Optics Lower "Coffin" 1 23.750 23.750Optics Bench to Bus Mounting Strut pairs 3 26.500 79.500Subtotal 237.54

STRAY LIGHT BAFFLE AssemblyAl Baffle Tube and Vane Assembly 1 112.8 141Primary (front) Solar Array 1 15.7 15.7Cooldown (rear) Solar Array 1 3.00 3SUBTOTAL 159.70Split (composite) Aperture Doors

Ribbed Composite Cover Halves 2 6.21 12.42Cover Hinge Mechanism 2 1.13 2.26Frangibolt Release Mechanisms 2 1.36 2.72

SUBTOTAL 17.40Total Mass 488.34

MASS (kg)

A.4.1.7 – CCD Chipset The RFP calls for a Earth observation system that is competitive with current observation systems.[A.4.1] This means that system must have a 1 pixel to 1 meter ratio. In the communications section, section 6, shows that the craft has a limited transmission rate. To compensate for the bottle-neck in the communication area of the spacecraft the amount of Earth image that can be seen at anyone time to about 10km diameter. Since each meter of image must have 1 pixel the system requires 10,000 pixels in diameter. Figure A.4.12 displays how 10,000 pixels are laid out.


Final Report 180

The current industry standard for pixels is 15 μm per pixel. This makes each 2000 x 2000 pixel chip 30 mm. 2000 x 2000 pixel chips were chosen for the design because that is the largest produced single CCD chip at this time [A.4.7]. This created the chipset as seen in Figure A.4.12.

Figure A.4.15: CCD Chipset

A.4.2 Space Viewing Telescope Calculating the diameter of the segmented Beryllium mirror, the equations for the sizing of a solid mirror were used at the suggestion of Prof. Finley of the department of physics at Purdue University. With a resolution of visible light (~5 micrometers) and a desired resolution of 0.05 arcsec, the following equation was used to find the diameter of the mirror: (A.4.9) While ideally the coefficient of 1.2 would be used, Hubble design teams used the factor of 1.4 for a more realistic estimate of size.[A.4.8] This provides the same diameter as Hubble, being 2.4 meters in diameter, since the same mission was to be accomplished by the SVT. Going into more detail about the design of the mirror, the focal length of the one-mirrored system was to be found with the following equations: Plate scale: (A.4.10)

Focal length: (A.4.11)

1.4resolutionDλ

≈ ⋅

deg2 12

(deg) 0.1667degcmCCDlength cm

resolution= =

180deg 687.5f platescale cmπ

= ⋅ =


Final Report 181

Radius of curvature of the mirror: 2 f = 13.74m (A.4.12) The focal length measures the distance from the base of the mirror to the CCD chip assembly that collects the images at the focal point of the mirror, held up by shape-memory NiTi struts, as described in the following section. CCD Support Design using GA Code Each leg of the tripod was designed as a tube to allow for the structure to carry the desired load while minimizing the mass. The minimum mass for Austenite was obtained using the Genetic Algorithm, or GA, optimization technique. The GA technique was chosen because of the multiple independent variables and the availability of a GA package provided by Professor William Crossley of the Purdue Aeronautics and Astronautics Engineering department. There were two primary constraints placed upon the structure yield and buckling. With these two constraints the mass function in equation A.4.13 was minimized within the GA code, where γ is the density, d is the diameter, t is the thickness, h is the height, and R is the base radius.

2 23m dt h Rγπ= + (A.4.13)

Required inputs: density of around 6.5 g/cm³ The design shown in Figure A.4.13 was used in the GA code with a varying diameter d and thickness t.

Figure A.4.13: CCD Support Structure[A.4.9]

The total mass of the CCD support structure is 12 kg, with a 9.65 mm diameter rod and a thickness of 3.326 mm. This supports a 2.26 kg load that can be applied by the CCD before failure would occur via buckling. This incorporates a safety factor of twice the max anticipated load.


Final Report 182

A.4.3 Cost To estimate the total cost of the SVT other than by SMAD equations, the individual components were priced and summed. The following table summarizes the sources used to price the components.

SVT Cost Estimate Beryllium Mirror[A.4.8] $13,564,800FGS[A.4.9] $30,000,000CCD chips[A.4.10] $5,000 NiTi CCD Supports[A.4.11] $9,302 Aluminum supports for mirror[A.4.12] $130 Total Cost of SVT $43,579,232

Table A.4.1: SVT cost estimate Segmented Beryllium Mirror - The segmented Beryllium mirror was based on a cost per square meter estimate made by a NASA study, comparing four different types of mirrors.[A.4.8]

Fine Guidance System - The Fine Guidance System estimate is based on the assumption that the Odyssey would employ rebuilt/reused FGS systems that had previously flown on other space missions. Not using brand new hardware dramatically decreasing the price of the system by tens of millions of dollars. [A.4.9]

CCD Chip Assembly - CCD chips, employed in most video recorders and cameras today, are relatively inexpensive. They can be found on average electronics websites for several hundred dollars usually. An overestimate was made for the CCD chip assembly since 1024 by 1024 pixeled CCD chips are not as common as 800 by 800 pixeled chips or others.[A.4.10]

NiTi Struts for CCD Chip Support - The estimate for the NiTi supports was made by finding the volume of the rods, using the density to find a weight, and then using the estimates from the Dynalloy website to price, based on weight. NiTi is currently sold in thin wire-form but can be specially ordered in thicker form.[A.4.11]

Aluminum Supports for Mirror - Simple Aluminum was used for the supports of the light-weight Beryllium mirror. Basic prices from current stock prices were assumed and applied to the weight obtained by using the GA optimization code. [A.4.12]


Final Report 183

A.4.4 References [A.4.1] Larson, Wiley J., and James R. Wertz, eds. Space Mission Analysis and Design. 7th ed.

El Segundo: Microcosm, Inc., 2005. [A.4.1] Weisshaar, Terrence. Purdue Advanced Research Projects Affiliates (PARPA)

REQUEST FOR PROPOSAL: Modular, Fractionated Space System Advanced Concept Technology Demonstrator." Purdue Advanced Research Projects Affiliates (PARPA) Mission Opportunity Statement. West Lafayette: Purdue University, 2006.

[A.4.3] Dewandre, Thierry M., Joachim J. Schulte-In-Den-Baumen, and Emmanuel Sein,

comps. Space Optics 1994: Space Instrumentation and Spacecraft Optics. SPIE: Bellingham, Wash, 1994.

[A.4.4] Born, Max, and Emil Wolf. Principles of Optics; Electromagnetic Theory of

Propagation, Interference and Diffraction of Light. 7th ed. Cambridge UP, 1999. [A.4.5] MatWeb, Your Source for Materials Information." Matweb: Material Property Data. 3

Dec. 2006 <http://www.matweb.com>. [A.4.6] Sun, C T. Mechanics of Aircraft Structures. New York: Whiley Interscience, 1998. [A.4.7] Barbe, D F. "Imaging with Charge Coupled Devices." AAIA (1990). [A.4.8] “Advanced Telescopes & Observatories Capability Roadmap Presentation to the NRC.”

NASA & APIO. 15 March 2005. [A.4.9] Jenkins, Ann.“Hubble Space Telescope Servicing: Demonstrating Space Logistics”

15 Nov. 2006. http://www.findarticles.com/p/articles/mi_qa3766/is_200501/ai_n15869331 [A.4.10] “Global Spec.” 20 Nov. 2006. http://GlobalSpec.com [A.4.11] “NiTi Price Guide.” 25 Nov 2006. http://www.dynalloy.com/PriceGuide.html [A.4.12]”Aluminum Pricing.” 30 Nov. 2006.

<http://www.moneyowl.co.uk/xal/EUR/forex.php?epi=XAL1_aluminium%20price&gclid=CIvv_LGi14gCFRJ5SQodnz7IOg>

http://www.findarticles.com/p/search?tb=art&qt=%22Jenkins%2C+Ann%22


Final Report 184

A.6 Communication and Data Handling A.6.1 – Link budget[A.6.1]

The link budget traces the power of the signal from transmission to reception for a given communication link. The link budget is a combination of inputs, estimations from charts, and calculated results. A.6.1.1 – Inputs The inputs into the link budget are explained below along with the variable representation: Frequency – f – The frequency is a property of the signal which was dictated by TDRS requirements. Power – P – The transmitter power in Watts. This is a function of the transceivers used and amount of available power. Transmitter line loss – Ll – The amount of power lost in the wiring of the communication hardware. This is an estimate that ranges from -1 dB to -3 dB. A loss of -2 dB was used in the design due to the short distance between the transceivers and antennas. Antenna diameter – D – TDRS’s antennas are fixed at 15 meters for the second generation satellites and 5 meters for the first generation ones. The 15 meter diameter was used in the design of the link budget as a worst-case scenario because a 5 meter results in a higher link margin which is discussed in section A6.1.2. The diameter of Odyssey’s antennas was a design variable that was varied to achieve the desired link margin. Antenna pointing error – e – The antenna pointing error is a combination of the attitude control accuracy of the satellite and accuracy of the antenna pointing system. An error of 0.1° was used as a conservative estimate of both systems. Actual pointing error was estimated at an order of magnitude less. Propagation path length – S – This is the distance in km that the signal has to travel from one antenna to the other. The estimate used for the link design was a conservative 50,000 km which is the approximate distance from TDRS to the center of the Earth plus 5,000 km. Propagation and polarization loss – La – Propagation and polarization loss is the loss in power caused by the signal passing through Earth’s atmosphere. There is no propagation and polarization loss in Odyssey’s communication links because both spacecraft are in orbit and it does not communicate directly with a ground station on the surface. Data rate – R – The data rate is the amount of information that can be transmitted in bits per second. Each link has a different maximum data rate dictated by TDRS. The Ku uplink data rate


Final Report 185

is not the maximum TDRS can provide, but there is no foreseeable need for an uplink speed greater than 1 Mbps. Bit-error rate (BER) – BER – The bit-error rate is the percentage of bits with errors divided by the total number of bits transmitted. A BER of 10-5 was used in the link budget because it is required for communication with TDRS System noise temperature – Ts – The system noise temperature is the power loss due to thermal interference in the antenna and hardware. The 2.7K temperature of space leads to low loss. An estimate of 10 dB was used in the link budget. Required Eb/No – Eb/No – E/N is the ratio of received energy per bit to noise density. It is a measure of quality for a communication system. The required E/N depends on the bit error rate and type of modulation and coding used. An E/N of 10 dB would provide the required bit error rate using TDRS compatible modulation schemes. Implementation loss – Limp – Implementation loss is a compensation factor in the link budget. It takes into consideration that hardware is not 100% efficient and there will be some power loss that is unaccounted for. A.6.1.2 – Calculations The following values are calculated by the spreadsheet and Matlab code from the inputs defined above. Transmitter power – P – The antenna power was provided in Watts and needs to be converted to dB. This is calculated using equation A.6.1.

10 log( )dB WP P= ⋅ (A.6.1) Antenna beamwidth – θ – The antenna beamwidth is the width in degrees of the beam that has half the power. The power of the signal decreases as distance from the centerline increases and the beamwidth is defined as two times the angle from the centerline to the line where the signal is at half power. The estimated antenna beamwidth is calculated using the antenna diameter and frequency in equation A.6.2 for both the transmitting antenna and receiving antenna.

21f D

θ =⋅

(A.6.2)

Peak antenna gain – Gpeak – The peak antenna gain is the maximum ratio of power radiated to the center of the coverage area to the power radiated by an isotropic antenna. The peak antenna gain for a circular antenna is calculated from the antenna beamwidth and assuming an efficiency of 55% by using equation A.6.3.

( )244.3 10 logpeakG θ= − ⋅ (A.6.3)


Final Report 186

Antenna pointing loss – Lθ – The antenna pointing loss is a function of the beamwidth and pointing error. This is the loss in signal power that results from the antennas not pointing directly where they are intended.

2

12 eLθ θ⎛ ⎞= − ⎜ ⎟⎝ ⎠

(A.6.4)

Antenna gain (net) – Gnet – The net antenna gain is defined as the peak antenna gain plus the pointing loss as shown in equation A.6.5.

net peakG G Lθ= + (A.6.5) Equiv. isotropic radiated power (EIRP) – EIRP – EIRP is the power that would have to be supplied to an isotropic antenna to achieve the same field strength at the receiver. The higher the gain of the antenna the higher the EIRP because there is more energy directed in a particular direction. Equation A.6.6 shows the calculation for the EIRP. The subscript t means it is the gain of the transmitting antenna.

l tEIRP P L G= + + (A.6.6)

Space loss – Ls – The space loss is the loss in power of the signal as it propagates through space. The loss is caused by energy dissipation as the waves travel and radiation interference from the space environment. The power loss is a function of the distance traveled and frequency of the signal. In equation A.6.7 c is the speed of light in m/s, S is the propagation length, and f is the frequency in Hz.

4scLSfπ

⎛ ⎞= ⎜ ⎟

⎝ ⎠ (A.6.7)

To calculate the space loss in dB use equation 6.8.

( ) ( ) ( ) ( )20 log 20 log 4 20 log 20 logsL c S fπ= ⋅ − ⋅ − ⋅ − ⋅ (A.6.8) Signal to noise ratio – Eb/No – The signal to noise ratio is the fundamental quantity that defines the quality of the signal when it is received. It is defined as the ratio of received energy per bit to noise density in a digital communication system. Any value greater than five is adequate for simple binary communication with a low probability of error. The higher the ratio, the better the signal and lower probability of error. Equation 6.9 provides the Eb/No in dB with all variables being defined in dB. The subscript t means it is the value for the transmitting antenna and the subscript r represents the receiving antenna.

( ) ( )228.6 10 log 10 logbl t pr s a r s

o

E P L G L L L G T RN

= − + + + + + + − ⋅ − ⋅ (A.6.9)


Final Report 187

Carrier to noise ratio – C/No – The carrier to noise ratio is defined as the power of the carrier signal to the power of the noise. Both the carrier to noise ratio and signal to noise ratio are measurements to the quality of the communication link. Equation 6.10 shows the relationship between the signal to noise ratio and carrier to noise ratio which involves the log of the data rate.

( )10 logb

o o

C E RN N

= + ⋅ (A.6.10)

Margin – M – The margin identifies if the specified link properties will work together to produce a viable link. The margin is calculated from the signal to noise ratio, the required signal to noise ratio, and the implementation loss. A margin of 3 dB will ensure a reliable communications link between the two satellites and a margin greater than 3 dB will provide an even more reliable link. Equation 6.11 shows how the link margin is calculated.

b bimp

o orequired

E EM LN N

= − + (A.6.11)

A.6.1.3 – Excel spreadsheet

Table A.6.1: Link budget spreadsheet Spacecraft Antenna Diamater [m] 0.5Spacecraft Pointing Accuracy 0.1

Item Symbol Units Source Uplink Downlink Uplink DownlinkTransmit antennaFrequency f GHz Input 13.775 15.003 2.1064 2.2875Transmitter Power P Watts Input 10.00 10.00 10.00 10.00

dB Calc 10.00 10.00 10.00 10.00Transmitter Line Loss Ll dB Input -2.00 -2.00 -2.00 -2.00Transmit Antenna Diameter Dt m Input 15.00 0.50 15.00 0.50Transmit Antenna Beamwidth θt deg Calc 9.84 0.36 5.00 0.05Peak Transmit Antenna Gain Gpt dBi Calc 24.44 53.24 30.32 69.58Transmit Antenna Pointing Error et deg Input 0.10 0.10 0.10 0.10Transmit Antenna Pointing Loss Lpt dB Calc 0.00 -0.94 0.00 -40.45Transmit Antenna Gain (net) Gt dBi Calc 24.44 52.30 30.32 29.12Equiv. Isotropic Radiated Power EIRP dBW Calc 32.44 60.30 38.32 37.12Propagation Path Length S km Input 50000.00 50000.00 50000.00 50000.00Space Loss Ls dB Calc -209.82 -210.56 -193.51 -194.23Propagation & Polarization Loss La dB Fig 13.10 0.00 0.00 0.00 0.00Receive antennaReceive Antenna Diameter Dr m Input 0.50 15.00 0.50 15.00Peak Receive Antenna Gain (net) Grp dBi Calc 34.57 64.85 18.25 48.51Receive Antenna Beamwidth θp deg Calc 3.05 0.09 19.94 0.61Receive Antenna Pointing Error er deg Input 0.10 0.10 0.10 0.10Receive Antenna Pointing Loss Lpr dB Calc -0.01 -13.78 0.00 -0.32Receive Antenna Gain Gr dBi Calc 34.55 51.07 18.25 48.19System Noise Temperature Ts K Table 13.10 10.00 10.00 10.00 10.00Data Rate R bps Input 1.049E+06 3.146E+08 3.072E+05 3.072E+05Signal-to-Noise Ratio Eb/No dB Calc 15.55 20.65 26.79 54.50Carrier-to-Noise Ratio C/No dB-Hz Calc 75.76 105.63 81.66 109.37Bit Error Rate BER Input 1.000E-05 1.000E-05 1.000E-05 1.000E-05Required E/N Req Eb/No dB Fig 13.9 10.00 10.00 10.00 10.00Implementation Loss dB Input -2.00 -2.00 -2.00 -2.00Margin dB Calc 3.55 8.65 14.79 42.50

InputEstimated from Graph


Final Report 188

A.6.1.4 – Matlab Below is a code written to perform the calculations in the Excel spreadsheet shown above. The code was written so ranges of variables could be tested at the same time. The code carries out the calculations and plots the results so the optimal solution can be identified. A.6.1.4.1 – Sample Matlab code %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %% %% Programmer: Matt Dennis %% Class: AAE 450 %% Date: 06 September 2006 %% %% Code Description: Communication link budget (downlink) %% %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% Inputs %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% % Link f = 13.2; % Frequency [GHz] rfl = -2; % Estimated RF loss [dB] S = 22000; % Propagation Length Ll = -1; % Transmitter Line Loss [db] % Transmitter Pw = 10; % Power input [Watts] Dt = 0.5; % Transmitter Diameter [m] et = .1; % Transmit Antenna Pointing Error [deg] % Reciever Dr = 15; % Recieve Antenna diameter [m] er = .1; % Recieve Antenna Pointing Error [deg] R = 300*1024^2; % Data Rate [bps] BER = 1e-5; % Bit Error Rate La = 0; % Propagation & Polarization Loss [dB] Ts = 2.7; % System Noise Temperature [K] REN = 10; % Required E/N [dB] Impl = -2; % Implementation Loss [dB] %%%%%%%%%%%%%%%%%%%%%%%%%%%%% Calculations %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% Ptdb = 10.*log10(Pw); tht = Dt.*f./21; Gpt = 44.3-10.*log10(tht.*tht); Lpt = -12.*(et./tht).^2; Gt = Lpt+Gpt; EIRP = Ptdb+Gt+Ll; Ls = 20.*LOG10(3.*10^8)-20.*LOG10(4.*pi)-20.*LOG10(S.*1000)-

20.*LOG10(f.*1024^3); Gpr = 20.*LOG10(pi)+20.*LOG10(Dr)+20*LOG10(f.*1024^3)+10.*LOG10(0.55)-

20.*LOG10(3*10^8); thr = 21./(f.*Dr);


Final Report 189

Lpr = -12.*(er./thr).^2; Gr = Lpr+Gpr; EN = Ptdb+Ll+Gt+Lpr+Ls+La+Gr+228.6-10.*log10(Ts)-10.*log10(R); CNo = EN+10.*log10(R); Mar = EN-REN+Impl; Dts = size(Dt); ers = size(er); Drs = size(Dr); if Dts(2) > 1 figure(1) plot(Dt,Mar,Dt,3,'-r') axis([min(Dt), max(Dt), -60, 80]) xlabel('Transmit Antenna Diameter [m]') ylabel('Margin [dB]') title('Downlink Margin vs Antenna Diameter (Ku-band)') legend('Link margin','Minimum margin allowed') end if Drs(2) > 1 figure(2) plot(Dr,Mar,Dr,3,'-r') axis([min(Dr), max(Dr), -60, 60]) xlabel('Recieve Antenna Diameter [m]') ylabel('Margin [dB] (Downlink)') title('Margin vs Antenna Diameter (Ku-band)') end if ers(2) > 1 figure(3) plot(er,Mar,er,3,'-r') xlabel('Transmit Antenna Pointing Error [deg]') ylabel('Margin [dB] (Downlink)') title('Margin vs Transmit Antenna Pointing Error (Ku-band)') end


Final Report 190

A.6.1.4.1 – Sample Matlab plots

Figure A.6.1 – Uplink Margin vs Antenna Diameter (S-band)

Figure A.6.2 – Downlink Margin vs. Antenna Diameter (S-band)

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9-60

-40

-20

0

20

40

60

80

Transmit Antenna Diameter [m]

Mar

gin

[dB

]

Downlink Margin vs Antenna Diameter (S-band)

Link marginMinimum margin allowed

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9-20

-10

0

10

20

30

40

Recieve Antenna Diameter [m]

Mar

gin

[dB

]

Uplink Margin vs Antenna Diameter (S-band)



Final Report 191

Figure A.6.3 – Uplink Margin vs. Antenna Diameter (Ku-band)

Figure A.6.4 - Downlink Margin vs. Antenna Diameter (Ku-band)

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9-60

-40

-20

0

20

40

60

80

Transmit Antenna Diameter [m]

Mar

gin

[dB

]

Downlink Margin vs Antenna Diameter (Ku-band)


0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9-20

-15

-10

-5

0

5

10

15

20

Recieve Antenna Diameter [m]

Mar

gin

[dB

]

Uplink Margin vs Antenna Diameter (Ku-band)



Final Report 192

A.6.2 – References [A.6.1] Larson, Wiley J., and James R. Wertz, eds. Space Mission Analysis and Design. 3rd ed.

El Segundo: Microcosm P, 2005


Final Report 193

A.7 – Structures and Mechanisms A.7.1 – Material Trade Study Seven materials were evaluated for their fitness for use in the primary structure. It was desirable to choose a material with the lowest thermal conductivity, the lowest thermal expansion coefficient, and the lowest structural mass for a given application. The material properties, and materials evaluated are shown below:

Table A.7.6: Properties of Evaluated Materials Tensile Compressive Young's Poisson's Thermal Thermal

Density Yield Yield Modulus Ratio Conductivity Expansion

kg/m^3 Mpa Mpa Gpa W/m-Kmicrom/m-C

Aluminum (2219-T851 Plate) 2850 414 320 72 0.33 130 23.6Steel (17-4PH H1150z Bar) 7860 634 660 201 0.29 15 10.8Magnesium (AZ31B H24 Sheet) 1170 150 165 45 0.35 96 26Titanium (Ti-6Al-4V Annealed Plate) 4430 828 855 110 0.342 6.6 9Thornel Graphite Cloth Composite 1400 103 NA 13 NAAluminum-Graphite MMC 2500 76.9 109 89 0 190 7.5Pyrolytic Graphite (GE Advanced Ceramics) 2180 80 100 20 0 300 0.5

Certain combinations of material properties for beams, columns, rods, and endplates will indicate which material choice will yield the least mass. These parameters are derived by relating the failure-condition equation to the mass of the member, and factoring out all properties intrinsic to the material.

Figure A.7.16: Simple member

For simple rods: Lhm 2*ρ= (A.7.11)

2hF

yield =σ FLmyield

*⎟⎟⎠

⎞⎜⎜⎝

⎛=

σρ

( A.7.12)


Final Report 194

For beam bending:

4*12/12/*

hhM

IMy

yield ==σ LMmyield

3/23/2 *

⎟⎟⎠

⎞⎜⎜⎝

⎛∝

σρ

(A.7.13)

For column buckling:

2

4

2

*L

chkEL

kEIFcr == 2* Lk

FE

m cr⎟⎠

⎞⎜⎝

⎛∝

ρ (A.7.14)

For plate failure, our equation for the stress on the plate, with constants derived for the maximum stress using the finite element analysis is as follows:

2

2

tPDCσσ = , tcDAtm ** 2ρρ == σσ

ρ PCDmcr

** 3

⎟⎟⎠

⎞⎜⎜⎝

⎛∝ ( A.7.15)

The materials were ranked according to combinations of these four properties to evaluate their fitness for each structural job:

Table A.7.7: Material Structural Performance Endplates Compressive Rod Tensile Rod Columns Beams

Fyρ

ρ/1

Fcy*

ρ/1

Fty*

Eρ * 3/2Fy

ρ

Aluminum (2219-T851 Plate) 159.32 8.91 6.88 335.88 60.92Steel (17-4PH H1150z Bar) 312.16 11.91 12.40 554.40 106.50Magnesium (AZ31B H24 Sheet) 95.53 7.09 7.80 174.41 41.44Titanium (Ti-6Al-4V Annealed Plate) 153.95 5.18 5.35 422.38 50.24Thornel Graphite Cloth Composite ~ ~ 13.59 388.29 ~ Aluminum-Graphite MMC 285.09 22.94 32.51 265.00 138.25Pyrolytic Graphite (GE Advanced Ceramics) 243.73 21.80 27.25 487.46 117.42

*From Peery[7.1], the others derived independently. Green is best, yellow is second The materials were ranked according to their performance in each structural capacity along with thermal expansion and thermal conductivity, each weighted equally. The rankings were subtracted from six and squared to give them a greater spread and give the overall best material prominence. The scoring tables are shown below:


Final Report 195

Table A.7.8: Ranking of Materials Thermal Properties Structural ApplicationsThermal Expansion

Thermal Conductivity Plate Beam Rod Column

Aluminum (2219-T851 Plate) 5 4 3 3 3 3Steel (17-4PH H1150z Bar) 4 2 6 4 4 6

Magnesium (AZ31B H24 Sheet) 6 3 1 1 2 1Titanium (Ti-6Al-4V Annealed Plate) 3 1 2 2 1 4Aluminum-Graphite MMC 2 5 4 5 6 2Pyrolytic Graphite (GE Advanced Ceramics) 1 6 5 6 5 5Weight 1 1 1

Table A.7.9: Scoring of Materials Score = (6-ranking)^2*weightPlate Beam Rod Column

Aluminum (2219-T851 Plate) 14 14 14 14Steel (17-4PH H1150z Bar) 20 24 24 20Magnesium (AZ31B H24 Sheet) 34 34 25 34Titanium (Ti-6Al-4V Annealed Plate) 50 50 59 38Aluminum-Graphite MMC 21 18 17 33Pyrolytic Graphite (GE Advanced Ceramics) 26 25 26 26

With consideration of the thermal properties of each material, titanium did the best in every category. Titanium was chosen as the primary structural material for the core sections and modules. A.7.2 – Finite Element Analysis Finite element analyses were conducted in ABAQUS to determine the behavior of more complicated pieces of our structure under loads. The primary assumption in this analysis was linear elasticity, valid for small deformations and stresses less than the yield stress[7.2]. Each situation was modeled, and the thickness, diameter, young’s modulus, load, and poison’s ratio were varied to solve the following equation for its exponents:

54321*, nnnnn PEvtDC=∂σ (A.7.16)

Table A.7.10: Variation of Parameters for Old Octagonal Endplate

D (m) t (mm) E (Gpa) v P (Pa) sig (Pa) delta (mm)1000 10 1.00E+10 0.3 100 6.92E+06 3.26E-011000 10 1.00E+10 0.3 1000 6.30E+07 3.59E+001000 10 1.00E+10 0.1 100 8.02E+06 3.83E-011000 10 1.00E+09 0.3 100 6.30E+06 3.59E+001000 5 1.00E+10 0.3 100 2.25E+07 2.38E+00

500 10 1.00E+10 0.3 100 1.80E+06 2.17E-02


Final Report 196

The equation was solved for variations on the old octagonal endplate, fixed at the boundaries, a section of skin attached to stringers with a concentrated center load, and an alignment fin with a concentrated edge load. The maximum displacements and stresses were recorded for each situation.

Table A.7.11: Stress and displacement compared to variation of skin thickness t (m) sigma (Pa) delta (m)

0.001 1.78E+06 9.31E-050.0015 8.09E+05 3.07E-050.0025 3.01E+05 8.94E-06

0.005 8.24E+04 2.55E-060.01 3.18E+04 9.33E-07

The model for the endplates derived above is as follows:

2

2

tPDCσσ = (A.7.17)

EtPDC3

4δδ = (A.7.8)

For the skin and alignment fins, the model is as follows:

2tFCσσ = (A.7.9)

EtFC

3δδ = (A.7.10)

Several boundary condition cases were evaluated, due to uncertainty in exactly how the endplates would be attached to the rest of the structure. The first case tested was to fix just the corners. Further cases were evaluated as the semester proceeded to obtain the plate constants for lighter versions of the plate. The final version of the endplate had holes cut in the sides to reduce mass, making it a bulkhead stiffener.


Final Report 197

Figure A.7.17: Two FEA cases analyzed

These plate constants, once obtained, were used to size the thickness of each component in the systems spreadsheet according to our model.

Table A.7.12: Obtaining the Plate Constants for each Situation Module EndplateBuilt in point corners -redone, but can't find the numbersMaterial E (Pa) v t (m) dR(m) sigma (Pa) delta (m) P (Pa) Csigma CdeltaTitanium 1.10E+11 0.342 0.01 1.5 3.83E+05 5.22E-06 1.00E+00 17.03 0.1135Built in 3cm cornersMaterial E (Pa) v t (m) dR(m) sigma (Pa) delta (m) P (Pa) Csigma CdeltaTitanium 1.10E+11 0.342 0.01 1.5 6.69E+04 2.17E-06 1.00E+00 2.97 0.0472Core Endplate2x2cm built in cornersMaterial E (Pa) v t (m) D (m) sigma (Pa) delta (m) P (Pa) Csigma CdeltaAluminum 7.20E+10 0.33 0.005 1 4.05E+04 1.70E-06 1.00E+00 1.01 0.0153New core plate with holes, 1.5cm built in cornersTitanium 1.10E+11 0.342 0.015 1 3.13E+03 1.11E-08 1.00E+00 0.70 0.0041

Load F on alignment fins, Long wrt widthMaterial E (Pa) v t (m) L (m) sigma (Pa) delta (m) F (N) Csigma CdeltaTitanium 1.10E+11 0.342 0.01 0.25 4.74E+04 1.04E-06 1 4.74E+00 1.04E-12

1.10E+11 0.342 0.005 0.25 1.87E+05 8.27E-06 1 4.67E+00 1.03E-121.10E+11 0.342 0.001 0.25 4.35E+06 1.02E-03 1 4.35E+00 1.02E-12

4.59E+00 1.03E-12


Final Report 198

Table A.7.13: Plate Constant Summary Boundary Condition with Octagonal Plate and Pressure Load P

Csigma CdeltaPinned Corners 6.303 0.0286Built In Corners 2.650 0.0283Built In Corner 2cm Regions 1.013 0.0153Pinned Edges 0.395 0.0130Built In Edges 0.164 0.0082New Plate Model 0.70 0.0041

Boundary Condition with Module Endplate and Pressure Load PCsigma Cdelta

Built in Corners 17.031 0.1135Built in Corners 3cm Regions 2.973 0.0472

Guidance Fin Csigma Cdelta4.586 1.03246E-12

Skin-Stringer Csigma Cdelta1.83 4.53

A.7.3 – Core Structure Design Two different core modules were sized according to the loads presented to each. Their members have similar definitions but different dimensions. The 2.3m core module of the experiment segment has eight modules attached to each face. The 5.5m core module of the optical segment has 15 one meter modules. These are attached to only three faces of the core, each stacked five high.

Figure A.7.18: Core primary structure


Final Report 199

Table A.7.14: Summary of Initial Core Properties Axial launch acceleration 68.6 m/sec^2Lateral launch acceleration 19.6 m/sec^2Maximum expected mass inside a module 489 kg

Optical Segment Core Experiment Segment CoreDiameter 1 m Diameter 1 mLength 5.5 m Length 2.3 mInternal Mass 730.297 kg Intenal Mass 626.38 kg(Mechanisms, piping, and wiring)

(Mechanisms, piping, and wiring)

Expected 15x 1m modules, five on three faces 7335 kg

Expected 8x 2m modules, one on each face 3912 kg

8x stringers 8x stringers4x buckling braces 16x torsion bars80x torsion bars Endplate area 0.16636 m^2Endplate area 0.16636 m^2

The core has five load bearing elements: The skin, stringers, endplates, torsion bars and alignment fins. Each of these components must be sized to handle failure modes pertaining to the shape. A summary of the considered failure modes is as follows:

Figure A.7.19: Failure modes considered

An impact load of 1000N was assumed to size the skin. The material chosen for all the basic structural components was titanium, which has a young’s modulus of 110E11 Pa, a yield stress of 828E8 Pa, and a density of 4430 kg/m^3. From our skin stress law, the load, and the yield stress, our skin has a thickness of 1.82mm and a maximum expected deflection of 1.2cm. The guidance fins were sized using the same equation models with their own plate constants.


Final Report 200

2/t

FCSF σσ = (A.7. 11)

EtFC

3δδ = (A.7. 12)

The endplates were sized to take the load of weight on top of or inside the module, modeled as a pressure distributed across the plate. The final design had no structure left on top, so the internal mass was assumed to load the plate.

2

2

/tPDC

SF σσ = (A.7.18)

Et

PDC3

4δδ = (A.7.14)

plate

axialernal

Agm

P int= (A.7.19)

The calculations yield 7.0=σC 0041.0=δC

The torsion bars were set to resist the torsion generated by axial acceleration acting on the module mass in a direction tangent to the core as shown in the diagram. The torsion bars were assumed to resist this load in tension. The torsion bars were located along each lateral face of the module. For the 5.5 m module, there are 5 lateral “faces” per size of the octagon taken between each buckling brace.

mgmrgmTorsion axialulestorsionaxialules 1*15 modmod == ∑ (A.7.16)

)sin(*8exp θcoreected r

TorsionF = 2/hFSFyield =σ (A.7.20)

The stringers were the most complex member to size. They were eventually made into parallelogram box-columns in order to provide the appropriate stiffness and compressive strength. Four failure modes acted on the stringers. The first two were stress due to axial compression from the axial force of gravity and tension or compression due to the core bending from lateral forces. The critical point of failure for these two modes was along the bottom of the vehicle where the stringers transfer their loads to the endplate and on to the propulsion segment. These modes were used to solve for the required cross sectional area.


Final Report 201

Figure A.7.20: Stringer cross section

The bending moment at the bottom of the core due to five sets of three modules at each level along the sums to 12.5m*3*modulemass*lateral acceleration.

core

lateralule

requiredc

axialulebendingaxialyield I

gmmA

gmSF

*3*5.128

15/ mod

sec,

modmax, +=+= σσσ (A.7.21)

)*)5.67sin(*)5.22(sin(*4 sec corecoretionccore rrAI += (A.7.22)

)*)5.67sin(*)5.22(sin(4*3*5.12

815

/ modmodsec,

corecoreyield

lateralule

yield

axialulebendingaxialrequiredc rr

gmmgmSFA

++=+=

σσσσ (A.7.20)

The next two failure modes were for buckling of the entire core and buckling of sections between the braces (or the endplates in the case of the experiment module). Buckling of the entire core turned out not to be significant with the cross sectional areas of the stringers already required.

2

2

kLEI

SFP corecr

π< (A.7.23)

Buckling between the braces was evaluated as if a stringer took one eighth of the compressive load for that interval of 1m and turned out to be more significant. This constraint was used to trade off between the dimensions of a and t in the stringer cross section diagram (width and thickness). In this manner, no increase in mass would be required, only an increase in the stringer width.

ESFPkL

I crrequired 2

2

π= (A.7.22)

)5.22(sin4 23

sec °= taI tionc (A.7.23)

atA tionc 4sec = (A.7.24)

)5.22(sin* 2sec °

=c

required

AI

a (A.7.25)


Final Report 202

In the case of the 2.3 meter module, the width of the stringers ended up becoming too wide with respect to the available room on the face, which has to accommodate the docking mechanism. An area multiplier was introduced. A thicker, more massive stringer cross section was used to constrain the width of the stringer to less than 3 cm.

Table A.7.15: Core dimensions, masses, and calculations Optical Segment Core Experiment Segment CoreInternal Mass Estimate 730 kg Internal Mass Estimate 626 kgSafety Factor 1.5 Safety Factor 1.5Radius 0.5 m Radius 0.5 mHeight 5.5 m Height 2 mEndplate Area 0.16636 m^2 Endplate Area 0.16636 m^2Endplate EndplateCsig 0.7 Csig 0.7Cdelta 0.0041 Cdelta 0.0041Pressure Load- 301144 Pa Pressure Load- 258293 PaThickness 0.0195 m Thickness 0.0181 mMass Of 1 Plate 14.402 kg Mass Of 1 Plate 13.338 kgMax Expected Deflection 0.0015 m Max Expected Deflection 0.0016 mStringer StringerMax Axial Load On Core 503442 N Max Axial Load On Core 268502 NMax Lateral Moment On Core 3.96E+05 Nm Max Lateral Moment On Core 7.67E+04 NmRequired Stringer Csectional Area 2.51E-04 m^2 Required Stringer Csectional Area 8.74E-05 m^2Inter-Brace Buckling Condition Inter-Brace Buckling ConditionCsectional Area 2.51E-04 m^2 Csectional Area 8.74E-05 m^2

Area Tuning Constant 2.00E+01Column Constant 0.50 Column Constant 0.50Required Moment Of Inertia 4.64E-08 m^4 Required Moment Of Inertia 4.95E-07 m^4a Dimension = 0.0178 m a Dimension = 0.0220 mThickness Dimension = 0.0035 m Thickness Dimension = 0.0199 mMass/Stringer 6.119 kg Mass/Stringer 15.486 kgSkin SkinMax Anticipated Docking Load 1000 N Max Anticipated Docking Load 1000 NCsigma 1.83 Csigma 1.83Cdelta 4.53 Cdelta 4.53Thickness 0.0018 m Thickness 0.0018 mMax Deflection Anticipated 0.0124 m Max Deflection Anticipated 0.0124 mMass 135.816 kg Mass 49.388 kgGuidance Fins Guidance FinsExpected Load 1000 N Expected Load 1000 NLength 0.1 m Length 0.25 mHeight 5 m Height 2 mCsigma 4.586 Csigma 4.586Thickness 0.002882355 m Thickness 0.002882355 mMass 6.384 kg Mass 6.384 kgTorsion Bars Torsion BarsExpected Torsion Radius 1 m Expected Torsion Radius 1 mExpected Torsion 143841 Nm Expected Torsion 9589 NmBar Angle Wrt Horizontal 69.06 deg Bar Angle Wrt Horizontal 69.06 degExpected Bar Load 38503 N Expected Bar Load 2567 NRequired Csection 6.97524E-05 m^2 Required Csection 4.65016E-06 m^2Dimension (Square Bar) 0.0084 m Dimension (Square Bar) 0.0022 mLength 2.80 m Length 2.80 mMass 0.865 kg Mass 0.058 kg


Final Report 203

Table A.7.16: Mass summary of core structure Optical Segment Core Mass Summary Experiment Segment Core Mass SummarySkin 135.82 kg Skin 49.39 kg8x Stringers 48.95 kg 8x Stringers 123.89 kg2x Endplates 28.80 kg 2x Endplates 26.68 kg4x Guidance Fins 25.54 kg 8x Guidance Fins 51.08 kg80x Torsion Bar 69.17 kg 16x Torsion Bar 0.92 kgMass of Core Structure 308.27 kg Mass of Core Structure 251.95 kg

A.7.4 – Module Structure Design The module was designed to attach to the core at the docking mechanism. The structure was designed to hang from the docking mechanism in tension. The internal equipment mass of the module was assumed to be mounted on the end-plate, applying uniform pressure to the plate. There were two versions of the module – a 2m long version, for use in the experiment segment and to house the earth viewing telescope, and a 1m long version for use everywhere else. The only difference between the two is the height of the module and the angle of the tension bars.

Figure A.7.21: Module loading diagram

The tension bars were sized to take the tension load from the edge of the endplate. The endplate is kept in static equilibrium by tension along the frame and tension bars, and a horizontal load where the endplate meets with the buckling brace or endplate of the core.

)sin(2062.0 2

θPmFt = (A.7.26)

PmFa21473.0= (A.7.27)

AP

yield =σ (A.7.28)


Final Report 204

The endplates were sized according to the plate equations with the plate constants for pressure-loaded module endplates.

2

2

/tPDC

SF σσ = (A.7.29)

Et

PDC3

4δδ = (A.7.30)

plate

axialernal

Agm

P int= (A.7.31)

The equations yield: 973.2=σC 01598.0=δC

Table A.7.17: Module dimensions, masses, and calculations

1m Module 2m ModuleExpected Internal Mass 489 kg Expected Internal Mass 489 kgExpected Launch Acceleration 68.6 m/sec^2 Expected Launch Acceleration 68.6 m/sec^2Safety Factor 1.5 Safety Factor 2Radius from Axis 0.5 m Radius from Axis 0.5 mRadial Length 1 m Radial Length 1 mHeight 1 m Height 1.5 mEndplate Area 0.7071 m^2 Area 0.7071 m^2Force Balance (Given Pressure P Across the Plate) Force Balance (Given Pressure P Across the Plate)Fb = 0.206 P*m^2 Fb = 0.206 P*m^2Fa = 0.147 P*m^2 Fa = 0.147 P*m^2Angle 26.6 deg Angle 56.3 degFt 0.461 Pm^2 Ft 0.248 Pm^2Endplate EndplatePressure 47465 Pa Pressure 47465 PaCsig 2.973 Csig 2.973thickness 0.0160 m thickness 0.0185 mmass 50.08 kg mass 57.83 kgStringers and Tension Bars Stringers and Tension BarsPressure 47465 Pa Pressure 47465 PaFa = 6992 N Fa = 6992 NFt = 21889 N Ft = 11765 NRequired Cross Section Required Cross SectionTension Bar 3.97E-05 m^2 Tension Bar 2.84E-05 m^2Stringer 1.27E-05 m^2 Stringer 1.69E-05 m^2Square Side Length Square Side LengthTension Bar 0.0063 m Tension Bar 0.0053 mStringer 0.0036 m Stringer 0.0041 mSkin SkinMax Anticipated Docking Load 1000 N Max Anticipated Docking Load 1000 NCsigma 1.83 Csigma 1.83Cdelta 4.53 Cdelta 4.53Thickness 0.0018 m Thickness 0.0021 mMax Deflection Anticipated 0.0124 m Max Deflection Anticipated 0.0093 mTotal Area 3.3066 m^2 Total Area 4.9598 m^2


Final Report 205

Table A.7.18: Module mass summary 1m Module Structural Mass Module Structural MassSkin 26.67 kg Skin 46.20 kg2x Tension Bars 0.50 kg 2x Tension Bars 0.45 kg4x Stringers 0.22 kg 4x Stringers 0.45 kgUpper Frame 0.20 kg Upper Frame 0.26 kgEnd Plate 50.08 kg End Plate 57.83 kgTotal 77.67 kg Total 105.19 kg

A.7.5 – Propulsion Segment Structural Configuration It was required to have 3 tanks, with a volume of 1.02 m3 per tank. The available options were the cylindrical and the spherical tanks. For the spherical tanks case, the diameter will be 0.8 m for each tank. For the cylindrical tanks, a parametric study was made to see the change of the length of the cylinder with its diameter for the same volume. Figure A.7.7 below shows this variation.

32

34

scc rlrVolume ππ += (A.7.32)

(a)


Final Report 206

(b)

Figure A.7.7 (a)Cylindrical tanks dimensions. (b)Change of the length of a cylindrical tank with its diameter for the same volume.

The marked point is an interesting point because it gives the option of stacking two tanks with a total diameter of approximately 1 m, but with a length of about 6 m, which is too long for a launch vehicle packaging. The option of using cylindrical tanks was discarded in favor of using spherical tanks. A.7.6 – Propulsion Segment Structural Design I. Analysis of the I-beams: There are three I-beams in the propulsion segment. The upper one is 3 m long and the other two are 1.48 each, all having the same cross-section dimensions. The same analysis can be made for the lower and upper beams. Below, the detailed analysis for the lower, 1.48 m-long I-beam is provided.

0.2 0.25 0.3 0.35 0.4 0.45 0.55

10

15

20

25

30

35

Required volume =61,907 in3 /tank =1.02 m3 /tank


Final Report 207

Step 1. Free body diagram and shear and moment diagrams.

Figure A.7.8 Free body diagram for the lower I-beams in the propulsion segment.

Payload = 9000 kg 7g*Payload/2 = 309 kN Hydrazine thank weight full of propellant = 7g*376 = 25.8 kN He thank weight full of pressurant = 7g*408 = 28 kN Reaction forces: Right force = A Left force = B

0 : 672yF A B kN Equation A.7.33= + =∑(309)(0.04) (26)()0.56 (28)(1.07) (1.44)(309) (1.48) 0AM B= + + + − =∑ (A.7.34)

339.1332.9

B kNA kN

==


Final Report 208

(a)

(b)

Figure A.7.9 (a) . FBD with the magnitudes of forces for the lower I-beams in the propulsion segment. (b)

Shear force diagram.

The maximum moment can be found to be approximately the area underneath the V(x) from x=0 m to x=0.52 m. M_max =25.9 kN.m.


Final Report 209

Step 2. Stress Analysis. Applying a bending stress equation for the chosen cross section dimensions:

MyI

σ = Equation A.7.35

Figure A.7.10 Dimensions of the I-beam cross section.

M=25.9 kN.m y=0.12 m I=0.0001556 m4 (about the neutral axis)

3625.9*10 *0.12 19.9*10

0.0001556My PaI

σ = = = (A.7.36)

This value is less than the ultimate tensile strength value for Aluminum 7075-T6 which equals 538 MPa. Applying the following equation for shear stress for the chosen cross section dimensions:

VQIb

τ = (A.7.37)

V=339.1 kN Q = 0.00081 m3 I=0.0001556 m4 (about the neutral axis) b=0.015 m (where the maximum shear stress is)

36(339.1*10 )(0.00081) 117.7*10

(0.0001556)(0.015)VQ PaIb

τ = = = (A.7.38)

This value is less than the ultimate shear stress value for Aluminum 7075-T6 which equals 331 MPa. For the upper, 3 m-long I-beam, the maximum moment equals 271 kN.m. The maximum bending stress equals 209 MPa, which is lower than the ultimate value. The maximum shear stress is the same as in the previous case.


Final Report 210

The weight for these three I-beams equals the weight of one 3 m-I-Beam and two 1.48 m-I-Beams:

125 2(61.7) 248.4Weight kg= + = . II. Analysis of the rods: The rods are sized for buckling and checked for ultimate stress. The critical load that will cause a rod to buckle shown below

2

2( )crEIP

kπ

= (A.7.39)

Applying this for the rods with k=1, l=1 m, d=0.06m, E=71 GPa, 4 4

2 9 2 92

62 2 2

(0.06)(71*10 )( ) (71*10 )( )64 64 0.45*10

( ) (1)(1) (1)(1)cr

dEIP N

k

π ππ ππ= = = = (A.7.40)

The actual load the rod will experience is about 309 kN, which is less than the critical value. Checking the stress across the cross sectional area:

FA

σ = (A.7.41)

3309*10 1.090.28

F MPaA

σ = = = (A.7.42)

This value is less than the ultimate tensile strength (538 MPa). The weight for these four rods equal to 31.4 kg:

234*(2780 )*(0.0028 )*(1 ) 31.4kgWeight m m kgm= = (A.7.43)

Thus, the final weight of the structure of the propulsion segment equals 280 kg. Additional weight is expected for secondary structures to making a total weight of about 350 kg. A.7.7 – Module Docking Design The design of the module docking mechanism is based off of the design in the paper Concepts and Technology Development for the Autonomous Assembly and Reconfiguration of Modular Space Systems by Lennon Rodgers and David W. Miller. The ideas in the paper by Rodger and Miller provided a starting point for the design of the module to segment and core to core docking mechanisms. The module to segment docking mechanism needed to solve three design challenges: provide a structural and hard-docking connection, provide a power interface, and allow for data transmission between the module and Odyssey’s computer system.


Final Report 211

The challenge of providing a structural and hard-dock connection was solved by the inclusion of the locking pins on the mechanism. There are two locking pins on each docking mechanism. In the design of the pins it was assumed that each pin will be carrying one half of the maximum load of a full module. In this case the maximum load was 94 kg for module material mass plus 489 kg, half the mass of the EVT. The plates used to construct the connection faces are 2 cm each located on the module’s bottom surface, the internal face of the connection on which the locking plate is located are mounted on the inside of the module. Between the two faces of the connector is a metal plate of the same thickness of the module material to provide a snug interface. When all of the plate thicknesses are added it results in the need for a pin length of over 6 cm. Since there are modules attached to the core at the launch of the Odyssey system the pins must provide enough structural support to survive launch and must analyzed with the appropriate forces. The pins were checked for yielding using a simple calculation. As this is a design concept, and if it were produced it would be subject to rigorous testing simple calculations suffice to produce the conceptual design. The equations used in pin sizing are shown as follows.

Cross sectional area of the pin

σσ FA

AF

=⇒= (A.7.44)

Using the area to obtain the radius

ππ ArrA =⇒= 2 (A.7.45)

The radius obtained was then doubled to account for the simplicity of calculation and to account for defects present in all materials and the liberal assumptions made in the design process. A pointed head was added to the pins to help align them in the holes which the pass through. The head of the pin also prevents it from sliding out of the connection once the slider plate moves into the locked position. When the pins are in the locked position it provides a structural and hard-docked connection completing the remaining interfaces. For the core to core mechanism the connector pins are doubled. The additional pins will cause the external plates and interfaces to be pushed together at four points. It helps to hold the connection tighter as it acts across a larger surface area with additional interfaces.


Final Report 212

Figure A.7.11: Module docking diagram

The interface for electrical power is provided by two metal tabs. The two tabs present on the connector will provide power for the module during its lifetime. The choice to incorporate two tabs on the connection was for redundancy and to ensure that the module is able to pull enough power through the connection. On the core to core connection the power tabs are doubled so that enough power is transferred across the connection to power the additional modules or subsystems. The data transfer interface is provided by the incorporation of a wireless network into the Odyssey’s computing and data handling system. The wireless system not only decreases the mass of the spacecraft but will also simplify the connection. When a wireless system is used the need for a hard connection is eliminated. A wireless transmitter/receiver, similar to a wireless card commonly found in laptop computers will be used. To allow the signal to pass through the inches of metal to the core of the satellite a hole is cut into the docking mechanism. This hole allows the signal to pass through for communication with the main system. The wireless network will be selected from the IEEE 802 family of wireless networking options. The technology for a fully wireless satellite system is not fully developed at the current time, but with the rate of testing in space and the pace of advancement in technology it is assumed that it will reach the proper level of maturity by the build date of Odyssey. The same concept is applied to core to core connections. The fuel interface supplied across the core to core connection required special considerations. This was done working with the propulsion subsystem team. The propulsion system uses valves to prevent the flow of the hydrazine or Helium across the unconnected interface. When a core to core docking procedure occurs the fuel interfaces will meet and be secured through a series of seals and secured using a screw down lock on the line interface. The lock will be detachable for the case when the core must be undocked.

Locking Plate

Connector Pin

Internal connection

face

External connection face

Wireless port

Power tab


Final Report 213

A.7.8 - References [A.7.1] Peery, David. Aircraft Structures. New York: McGraw-Hill, 1950. 279. [A.7.2] Cook, Robert D., David S. Malkus, Michael E. Plesha, and Robert J. Witt. Concepts and

Applications of Finite Element Analysis. 4th ed. Madison: John Wiley & Sons, 2002. 19.


Final Report 214

A.8 – Attitude Control System (ACS) A.8.1 - ∆v Budget The launch vehicle provides an orbit insertion accuracy of ±10 km altitude and ±0.1° inclination. An estimate of the ∆v required to correct the orientation can be computed using a Hohmann transfer. For a two-impulse maneuver, a Hohmann transfer provides the minimum ∆v. First, a ∆v is applied tangentially to the velocity vector of the original orbit to get on to the transfer ellipse at its perigee. Next, another ∆v is applied tangentially to the velocity vector of the transfer ellipse at apogee to circularize the orbit. This can be seen in Figure 1.[A.8.1]

Figure 1: Hohmann transfer[A.8.1]

In order to determine the total ∆v required for the transfer, the initial and final radii, as well as the initial and final velocities must be determined. The radius of the initial orbit (r1) can be determined according to Equation 1. In this equation, the semi-major axis (a) is 7163.19 km. Equation 2 simply states that that radius of the final orbit (r2) is simply equal to the semi-major axis.

kmar 101 ±= (A.8.1)

ar =2 (A.8.2)


Final Report 215

The velocities can be determined from Equation 3. This simplified equation takes advantage of the fact that the orbits are circular. The gravitational parameter of the Earth is μ = 398600.4418 km3/s2.

rvc

μ= (A.8.3)

Next, the semi-major axis and the eccentricity of the transfer ellipse can be calculated from Equations 4 and 5.

( )21 rraT += (A.8.4)

TT a

re 11−= (A.8.5)

Given these values, the velocities and perigee and apogee on the transfer ellipse can be calculated from Equations 6 and 7.

⎟⎟⎠

⎞⎜⎜⎝

⎛−=

Tp ar

vT

12

1

μ (A.8.6)

⎟⎟⎠

⎞⎜⎜⎝

⎛−=

Ta ar

vT

12

2

μ (A.8.7)

Now that the velocities have been determined, the Δv’s required for the Hohmann transfer are given in Equations 8 and 9.

11 vvv

Tp −=Δ (A.8.8)

22 vvvTa −=Δ (A.8.9)

Finally, the total Δv required for the Hohmann transfer is found using Equation 10.

21 vvvTOTAL Δ+Δ=Δ (A.8.10)

Using the method described here, it was calculated that for the maximum insertion error (10 km), it would require approximately 5.2 m/s of Δv for a corrective adjustment. Note: Equations 3-10 are taken directly from AAE 532 lecture notes. See Reference 1 for more details. Figure 2 shows the Δv requirement for insertion errors from zero to ten kilometers.


Final Report 216

1 2 3 4 5 6 7 8 9 100.5

1

1.5

2

2.5

3

3.5

4

4.5

5

5.5

Orbit insertion error [km]

Req

uire

d Δ

v fo

r cor

rect

ion

[m/s

]

Required Δ v to correction orbit insertion errors

Figure 2: Required ∆v vs. altitude orbit insertion error

This corrects the orbit insertion error in altitude. The required Δv to correct the orbit insertion error in inclination can be determined from Equation 11.[A.8.2]

⎟⎠⎞

⎜⎝⎛ Δ

=Δ2

sin2 ii

vvv (A.8.11)

The initial velocity (vi) of the vehicle is 7.46 km/s. Then the Δv for the maximum inclination error (0.1°) is 13.5 m/s. Then the maximum required ∆v to correct the altitude and inclination errors introduced by the launch vehicle is 18.7 m/s to be provided by the propulsion segment of the Odyssey. It was also necessary to calculate the ∆v required to compensate for atmospheric drag, the “principal nongravitational force acting on satellites in low-Earth orbit.”[A.8.2] This value can be determined from Equation 12. In this equation, CD is the drag coefficient, A is the satellite’s cross-sectional area, m is the satellite’s mass, ρ is atmospheric density, a, is the semi-major axis of the orbit.[A.8.2]

aVm

ACv Drev ρπ ⎟

⎠⎞

⎜⎝⎛=Δ (A.8.12)


Final Report 217

The drag coefficient, CD, is estimated to be 2.2. The cross-sectional area, A, is estimated to be 100 m2. The final satellite mass, m, is 9487 kg. The atmospheric density, ρ, was estimated at 8.37x10-14 kg/m3. Finally, as stated above, the semi-major axis, a, is 7163.19 km. Using these values, it was determined that the propulsion section would have to provide approximately 1 (m/s)/yr of ∆v to maintain the orbit. Because the atmospheric density and solar activity vary greatly, conservative values were chosen to estimate the ∆v. It is required that the satellite carry enough fuel aboard to deorbit at any time. From the Earth Satellite Parameters table in Space Mission Analysis and Design, it was determined that at the given altitude, the propulsion segment would have to provide 208.3 m/s of Δv to deorbit the satellite.[2] Finally, the Δv required to compensate for other environmental torques was estimated at approximately 5 (m/s)/yr. The variables in the torque equations have a high degree of unpredictability; many also require more system definition to determine. For this reason, a conservative estimate was chosen to account for these torques. A.8.2 - Center of mass The center of mass of the system was found using Equations 13-15. In these equations M is the total mass of the system (9487 kg). The mass of the individual components is mi. V1i, V2i, and V3i are the locations of the individual components measured from the end of the optical segment that connect to the propulsion segment. The center of mass of the Odyssey was found to be CM = -2.51 V1 + 0.03 V3 [m].

∑=

=n

iiiCM

VmM

V1

111 (A.8.13)

∑=

=n

iiiCM

VmM

V1

221 (A.8.14)

∑=

=n

iiiCM

VmM

V1

331 (A.8.15)

Here is the spreadsheet used to determine the final value of the CM. Module Location* Mass V1 (m) V2 (m) V3 (m) Module (kg) Payload (kg) V1*M (m-kg) V2*M (m-kg) V3*M (m-kg)1.1 -0.5000 -0.7071 -0.7071 77.67 12 -44.835 -63.406 -63.406 1.2 -1.5000 -0.7071 -0.7071 77.67 34 -167.505 -78.963 -78.963 1.3 -2.5000 -0.7071 -0.7071 77.67 12 -224.175 -63.406 -63.406 1.4 -3.5000 -0.7071 -0.7071 77.67 135 -744.345 -150.380 -150.380 1.5 -4.5000 -0.7071 -0.7071 77.67 50 -574.515 -90.276 -90.276 2.1 -0.5000 0.0000 -1.0000 77.67 12 -44.835 0.000 -89.670


Final Report 218

2.2 -1.5000 0.0000 -1.0000 77.67 244.5 -483.255 0.000 -322.170 2.3 -2.5000 0.0000 -1.0000 77.67 244.5 -805.425 0.000 -322.170 2.4 -3.5000 0.0000 -1.0000 77.67 0 -271.845 0.000 -77.670 2.5 -4.5000 0.0000 -1.0000 77.67 135 -957.015 0.000 -212.670 3.1 -0.5000 0.7071 -0.7071 77.67 12 -44.835 63.406 -63.406 3.2 -1.5000 0.7071 -0.7071 77.67 34 -167.505 78.963 -78.963 3.3 -2.5000 0.7071 -0.7071 77.67 12 -224.175 63.406 -63.406 3.4 -3.5000 0.7071 -0.7071 77.67 135 -744.345 150.380 -150.380 3.5 -4.5000 0.7071 -0.7071 77.67 50 -574.515 90.276 -90.276 2287.05 -2.6554 0.0000 -0.8383 *Assumes mass located in center of each module Super Module Location Mass V1 V2 V3 V1*M V2*M V3*M Module mass -2.5 0 0.75 2000.00 -5000.00 0.00 1500.00 FGS1 -2.5 -1 0.75 220 -550.00 -220 165.00 FGS2 -2.5 0 0.75 220 -550.00 0 165.00 FGS3 -2.5 1 0.75 220 -550.00 220 165.00 Space mirror -2.5 0 0.75 223 -557.50 0 167.25 TOTAL -2.5 0 0.75 2883.00 Core Location Mass V1 V2 V3 V1*M V2*M V3*M Core Mass -2.5 0 0 369 -922.50 0.00 0.00 Other1 0.00 0.00 0.00 Other2 0.00 0.00 0.00 TOTAL -2.5 0 0 369 Center of mass Mass (kg) V1 V2 V3 V1*M V2*M V3*M Super module -2.5000 0.0000 0.7500 2883.00 -7207.50 0.00 2162.25Modules 1,2,3 -2.6554 0.0000 -0.8383 2287.05 -6073.13 0.00 -1917.21Core -2.5000 0.0000 0.0000 369.00 -922.50 0.00 0.00 Optical segment -2.5642 0.0000 0.0442 5539.05 -14203.13 0.00 245.04Propulsion segment 0.7500 0.0000 0.0000 2000.00 1500.00 0.00 0.00Experiment segment -6.1500 0.0000 0.0000 1719.89 -10577.32 0.00 0.00 Center of mass of system -2.51 0.00 0.03


Final Report 219

MATLAB code constants.m %Constants global r_earth mu_earth r_earth = 6378.14; %Earth equatorial radius, km mu_earth = 398600.441800; %Earth's gravitational parameter, km^3/s^2 insertion_error.m %Calculates delta-v required to correct orbit %insertion error of the launch vehicle clear clc global mu_earth %km^3/s^2 r2 = 7163.19; %km v2 = sqrt(mu_earth/r2); %km/s for i = 1:10 x(i) = i; %km r1 = r2 - i; %km v1 = sqrt(mu_earth/r1); %km/s aT = 0.5*(r1+r2); %km eT = 1 - r1/aT; vpT = sqrt(mu_earth*(2/r1-1/aT)); %km/s dv1 = vpT - v1; %km/s vaT = sqrt(mu_earth*(2/r2-1/aT)); %km/s dv2 = v2 - vaT; %km/s dv_TOT(i) = dv1 + dv2; %km/s end plot(x,dv_TOT*1000) grid on xlabel('Orbit insertion error [km]') ylabel('Required \Delta v for correction [m/s]') title('Required \Delta v to correction orbit insertion errors') drag.m %Calculation of delta-v required %to compensate for atmospheric drag clear clc global r_earth


Final Report 220

rho = 8.37e-14; %atmospheric density, kg/m^3 Cd = 2.2; %drag coefficient m = 9487; %satellite mass, kg A = 5*3; %cross-sectional area, m^2 V = 7.47*1000; %velocity of spacecraft, m/s a = (r_earth+785)*1000; %semi-major axis, m R = 358; %revs/cycle D = 25; %days/cycle s = 358*365/25; a_D = -(1/2)*rho*(Cd*A/m)*V^2; %m/s^2 dv_drag_rev = pi*(Cd*A/m)*rho*a*V; %(m/s)/rev dv_drag_yr = s*dv_drag_rev %(m/s)/year


Final Report 221

A.8.3 - References [A.8.1] Howell, K.C. “Orbital Mechanics.” AAE 532 Lecture Notes. Purdue University, West

Lafayette, IN. Fall 2006. [A.8.2] Wertz, James and Wiley Larson. Space Mission Analysis and Design.

Microcosm/Kluwer. 1999.


Final Report 222

A.9 Launch Vehicle A.9.1 – Reliability The Falcon launch series is designed around cost and reliability. One of the biggest assets to creating a reliable design is studying historical data, and determine were the majority of failures occur. SpaceX performed its own study as well as looking at studies from other organizations such as the Aerospace Corporation. The two tables below show where the majority of failures occurred during the 90’s and early 2000’s.

Table A.9.1: U.S. Launch Failures from 1984 – 2004[A.9.1]

Table A.9.2 Aerospace Corporation Launch Vehicle Failure Study[A.9.1]

Based on these studies SpaceX determined that propulsion, avionics, and separation events were the biggest cause of launch vehicle failure in the modern space age. From this notion SpaceX designed its Falcon launch vehicles with redundant systems in these areas. The propulsion system is a cluster of Merlin engines using new and proven engine components. Some of these components come from the Saturn V engines which had a 100% mission success rate during


Final Report 223

Apollo. Not only are the components reliable, but since it is a cluster design any single engine failure can be compensated by the remaining four or eight engines, depending on the model. Propellant is fed via a single shaft, dual impeller turbo-pump operating on a gas generator cycle. The turbo-pump also provides the high pressure kerosene for the hydraulic actuators, which then recycles into the low pressure inlet. This eliminates the need for a separate hydraulic power system and means that thrust vector control failure by running out of hydraulic fluid is not possible. A third use of the turbo-pump is to provide roll control by actuating the turbine exhaust nozzle[A.9.2]. The 304s impulse makes the Merlin the best performing gas generator cycle kerosene engine built.

Figure A.9.1 Merlin Engine[A.9.2]

The avionics system is a triple redundant system with inertial navigation, and GPS for increased insertion accuracy. The insertion accuracies of these launch vehicles are just as accurate and in some instances better than the most expensive current vehicles. Along with the avionics are the separation devices which use dual initiated fully space qualified separation bolts. These bolts have a zero failure record in other launch vehicles. Thus based on historical launch data, and combining new and old design components. SpaceX has taken the best of the old and new launch vehicles, creating the most reliable system. Figure A.9.2 shows the predicted reliability of SpaceX vehicles.


Final Report 224

Figure 2: Expected Failure Rates All Causes Based on Historical Average[A.9.1]

A.9.2 - Cost Along with the design reducing much of the cost of the systems the operation of the SpaceX Corporation reduces the cost. SpaceX uses a flat management system to reduce cost. Organic organizations have a flat structure with only one or two levels of management. Flat organizations emphasize a decentralized approach to management that encourages high employee involvement in decisions. The purpose of this structure is to create independent small businesses or enterprises that can rapidly respond to customers' needs or changes in the business environment. The supervisor tends to have a more personal relationship with employees [1][A.9.3]. Along with the flat organization the use of an animated launch countdown reduces cost on personnel. Finally, SpaceX only produces one product, the Falcon launch vehicles. This allows all recourses and personnel to focus on one project creating a more organized effort with fewer managerial positions to communicate with other divisions.


Final Report 225

A.9.3 - References [A.9.1] "Space Launch Vehicle Reliability." 18 Oct. 2004.

The Aerospace Corporation. 21 Sept. 2006 <http://www.spacex.com/>. [A.9.2] "Falcon Engines." 2003. 12 Oct. 2006 <www.spaceX.com>. [A.9.3] Allen, Gemmy. "Management Modern." Organizing Process. 1998. 4 Dec. 2006

<http://telecollege.dcccd.edu/mgmt1374/book_contents/3organizing/org_process/org_process.htm>


Final Report 226

A.10 – Propulsion Appendix A.10.1 – Sizing of the Altitude Maintenance Subsystem Sizing for the altitude maintenance subsystem was more of a qualitative process than a quantitative one. From class notes for AAE 590E – introduction to electric propulsion it was established that current arcjet technology allows for a maximum specific impulse of 420 sec for hydrazine arcjets in the one kilo-watt range[A.10.1]. A mass to impulse correlation of 4 KG/kw for an arcjet with an Isp of 450 sec was found on page 704 of Space Mission Analysis and Design[A.10.2]. Using this correlation the thruster’s mass was estimated at 7 kg, including fittings and power conditioning electronics. The size envelope for the thruster was set at 10 cm x 10 cm x 50 cm using sizing information for a 1.8 kw unit produced by Primex[A.10.3]. The performance characterizes of the thruster were determined using the analytical performance data for arcjets on page 558 of Space Propulsion Analysis and Design[A.10.4]. The desired Isp and power requirements corresponded to a thrust performance of 0.17 N. A Δv requirement of 1 m/s/yr was supplied by the orbits group. Using equation 1 it was determined that 2.42 kg of propellant were required to supply the necessary Δv for orbit maintenance.

⎟⎠

⎞⎜⎝

⎛ −=Δ−

spoIgV

spacecraftpropellant eMM 1 (A.10.1)

The systems team decided that a maximum of 24 hours could be devoted annually to reboosting maneuvers. This stemmed from the operational requirement for an annual 24 shutdown of imaging operations to accommodate software upgrades. To satisfy this requirement it was necessary to prove that the burn time of the arcjet would be less than 24 hours. Equation 2 was used to calculate a burn time of 16.28 hours for the determined propellant weight, Isp, and thrust4.

FgIM

t sppropellantburn

0= (A.10.2)

In addition to altitude maintenance, the arcjet can be used to deorbit the spacecraft. Due to the large mass of the Odyssey it was decided that enough propellant should be held in reserve to perform a controlled deorbit of the spacecraft at any time. The Δv required for this maneuver was calculated by the orbits team to be 208 m/s. The propellant mass required for deorbiting was calculated to be 464.82 kg using equation 1.

A.10.2 – Attitude Control Thruster Sizing It was decided by the propulsion and ACS groups that the attitude control thruster size should remain in the 40-50 N range. This decision was not entirely arbitrary. The Mars Exploration


Final Report 227

Rover delivery vehicle used 4.4 N thrusters and had a total vehicle mass of 1062 kg[A.10.5]. Scaling this up to a 10,000 kg vehicle yields 45 N thrusters. This corresponded with known mass data for hydrazine thrusters produced by Marquardt and Aerojet. The Aerojet units were selected due to their low mass (3 kg per thruster) and high Isp (220 sec)2[A.10.2]. The positioning and arrangement of the thrusters was determined though the process outlined in section 10.4. The center of mass of the Odyssey was estimated by the systems team, and used by the propulsion team to estimate thruster moments. Using the center of mass approximation it was determined that the pitch and yaw thrusters have a moment arm of about 5 m and that the roll thrusters have a moment arm of about 1.8 m. System requirement 1.5 mandated that the attitude control thrusters be capable of rotating the spacecraft faster than the reaction wheels can. The reaction wheels can rotate the spacecraft 90o in 14 minutes (0.001869 rad/sec). It was decided that the system should carry enough propellant to accelerate the vehicle up to 3.14 rad/sec around the yaw and pitch axes per refueling period. The same amount of propellant would be allotted for roll acceleration. It is important to note that the spacecraft would never be rotated at these velocities, but that they represent the total angular Δv allotted for maneuvering. Since the thrusters are canted 45o off axis, the force supplied normal to the direction of rotation by each is only 31.8 N (~10 lbf). A moment of inertia estimate for the vehicle of 13,557.34 kg-m2 (10,000 slug-ft2) was obtained from SMAD[A.10.2]. The total thruster burn time for rotational maneuvering was calculated using equation 3 using the known values of thrust, Isp, radius, and angular velocity6. These yielded 157 seconds of pitch and 314 seconds of yaw thruster firing per refueling period. The thruster configuration dictates that four simultaneously firing thrusters are required per pitch maneuver. Two simultaneously firing thrusters are required for yaw maneuvers.

nTrI

t vb

ω= (A.10.3)

The propellant consumed during the burn was calculated using equation 4[A.10.6]. This yielded a propellant requirement of 85.64 kg of propellant for pitch and yaw maneuvering. Since the same amount of propellant was allotted for roll maneuvering, the process was repeated in reverse to yield a roll maneuver burn time and angular velocity of 157 sec and 1.13 rad/sec respectively.

sp

bp I

nTtw = (A.10.4)

According to the ACS group it is necessary to unload the reaction wheels once every 4 days (92 times a year). The maximum momentum of the chosen wheels is 265 kg-m-s. The propellant needed to unload each wheel was then calculated using equation 5. 2.55 kg of propellant are required annually to unload each reaction wheel.


Final Report 228

sp

wp LI

IW = (A.10.5)

A.10.3 – Fuel Distribution Subsystem Tank Sizing Figure A.10.1 shows a breakdown of the Odyssey propellant budget. The propulsion team decided that the spacecraft should carry a 12 month ACS propellant reserve in case of a delay in refueling by Orbital Express. This dictated a minimum hydrazine storage capacity of 998.78 kg.

Hydrazine Mass (kg)

Altitude maintenance 2.42Reaction wheel unloading 7.66

Attitude control maneuvering 264.57Total annual ACS and AMS requirement 274.65

Deorbit requirement 464.82

Total hydrazine requirement 739.47

FDS Capacity 998.78Reserve 274.65

Figure A.10.1 – Annual Hydrazine Budget Breakdown

The decision was made to utilize three propellant tanks. With this configuration a single tank malfunction would result in the loss of 333 kg of propellant. This would eliminate the emergency reserve and part of the de-orbit supply (though enough would remain for a semi-controlled de-orbit). A two tank failure would eliminate both the emergency supply and orbit reserve. Neither failure would affect the spacecraft’s operational propellant budget. A Four tank configuration was considered but the additional tank, plumbing, and valve masses made it impractical. Sizing of the hydrazine tanks was completed using the sizing process from the AIAA Element of Spacecraft Design home study correspondence course material[A.10.6]. An outline of the process is shown below. Calculate:

1. Propellant volume using the hydrazine mass requirement and the known hydrazine density of 1,017.03 kg/m3

2. Unusable propellant volume (3% of propellant volume)

3. Initial ullage volume using 3

unusableusableullage

VVV =

4. Tank volume before provisions for a bladder by summing the three volumes

5. Radius of the tank bladder using 3 tan*75.π

kVr =

6. Area of the bladder using Ab=2πr2 7. Bladder volume using a standard elastomer thickness of 1.91 mm


Final Report 229

8. Required tank volume using Vreq=Vtank+Vbladder The next step in the process is to calculate the structural mass of the tank. This involved determining the operating pressure for the propellant tanks (ullage pressure). For this application the ullage pressure was selected to be 5.52 MPa (800 psi). This selection was based on a desired propellant pressure at the thruster inlet of roughly 3.79 MPa (550 psi). This is typical for a thruster with a chamber pressure in the 2.41 MPa (350 psi) range. A 15% pressure loss for the fuel lines was estimated using a pressure drop calculator. Another 15% of pressure loss was added to account for the flow across the thruster venturi (flow metering device requiring a 15% pressure gradient). This yielded a tank ullage pressure of 5.52 MPa. Structural mass was calculated using the process below[A.10.6]. The hydrazine tanks are constructed from titanium resulting in an allowable stress of 553.41 MPa (safety factor of 1.5 included). Each tank is fitted with an internal elastomer bladder per industry standard. The bladders are necessary due to the behavioral nature of liquid-gas fluid interaction in microgravity. In short, the Helium will diffuse into the hydrazine and form bubbles. The bladder is used to physically separate the two and insure proper pressurization[A.10.6]. Calculate:

1. Pressure force on the tank walls using F=Pπr2

2. Stress on the wall using σ2

Pr=t

3. Weight of elastomer tank bladder using ( )( )33

34 rtrM bladder −+= πρ

4. Tank mass using ( ) bladdert MrtrM −−+= 33)(34 πρ

This process is for sizing spherical tanks. It yielded a tank outer diameter of 0.96 m and a tank mass of 35.50 kg. With this diameter the tanks could be arranged in a triangular pattern within the 2.8 m diameter of the propulsion segment. The spreadsheet was modified to calculate the mass of cylindrical tanks as well. Cylindrical tanks are not as structurally efficient as spherical tanks. The modification was done in case the diameter of the propulsion segment needed to be reduced. However, the propulsion segment diameter never changed and the cylindrical sizing calculator was never used. Due to its length and complexity its details will not be included here. Helium tanks were sized using the same approach, but without including bladders. Three tanks were chosen for redundancy and to minimize the amount of manifolding needed to connect to the three hydrazine tanks. The tanks are also made of titanium. A maximum operating pressure of 55.20 Mpa was specified per industry standard. It was decided that each tank should contain enough Helium to fully pressurize and empty a hydrazine tank 4 times. This was done for two reasons: redundancy and expandability. The system has a high level of redundancy as the spacecraft can operate off of a single Helium tank


Final Report 230

for up to two years. Furthermore, the extra Helium carried can be use to pressurize future experiments and refuel Helium cold gas thruster systems. The mass of Helium required to fully pressurize one hydrazine tank was calculated using the Ideal Gas Law with 5.52 MPa as the pressure value, the hydrazine tank volume sans bladder as the volume value, and 20o C (system operating temperature) as the temperature value. This yielded a mass of 9.01 kg per pressurization cycle. Each tank can perform four pressurizations yielding a total Helium mass of 108.15 kg divided among 3 tanks. The Helium volume per tank was calculated using the Ideal Gas Law with 55.20 MPa for the pressure value. Structural weight was calculated with the same method used for the hydrazine tanks. It yielded a Helium tank diameter of 0.51 m and a tank mass of 41.89 kg. A.10.4 – Fuel Distribution System Plumbing Layout and Cost The plumbing layout of the Odyssey was designed to allow multiple out-of-nominal valve failures per fuel line without affecting mission operations. The most likely sources of failure are valves leaking or jamming in the open position. For these cases redundancy was assured by having multiple isolation valves per fuel line. For example: Assume that the poppet valve in a propulsion segment yaw thruster were to jam in the open position. The thruster could be shutoff by closing the thruster isolation valve located on the fuel feed to that thruster. This valve would then be actuated to control the thruster’s firing. If that valve were to leak, the cluster isolation valve located on the cluster feed line would be used to isolate the leak. The two propulsion segment ACS fuel line isolation valves could be used to isolate the cluster if the cluster isolation valve were to leak or fail open. A three tiered failure like this is highly improbable, yet perfectly within the fault tolerance of the system. Another example involves an optics segment roll thruster isolation valve being stuck in the closed position, disabling the thruster. In this case the opposite set of roll thrusters would be used to roll the spacecraft in the opposite direction for all roll maneuvers. This may increase the time it takes to roll to an orientation, but the failure does not kill the system. Multiple thruster failures can be counteracted by varying the pulsing duration and frequency of other thrusters. Since each thruster is canted 45o off-axis firing one affects motion in two directions. Any alternate maneuvering will result in greater fuel consumption than nominal maneuvers, but will not reduce spacecraft maneuverability. The Odyssey plumbing and instrumentation diagram is shown in figure A.10.2a. The bill of materials for the plumbing configuration is shown below. All costs are educated estimates based on pricing experiences and expert input. The vendors reflect the industry leaders for the specific components.


Final Report 231

Table A.10.1 – Odyssey Propulsion Bill of Materials

Component Quantity Vendor Unit Cost Total Cost

Hydrazine Tank 3 ATK Space Operations $75,000 $225,000

Helium Tank 3 ATK Space Operations $100,000 $300,000

45 N Monoprop Thruster 12 Aerojet $200,000 $2,400,000

170 mN Arcjet 1 Primex $1,000,000 $1,000,000

Pneumatic Ball Valve 23 Moog or Marotta $10,000 $230,000

Solenoid Poppet Valve 29 Moog or Marotta $3,000 $87,000

Pressure Regulator 7 Grove $15,000 $105,000

Check Valve 17 Moog or Marotta $150 $2,550

Relief Valve 6 Moog or Marotta $1,500 $9,000

Burst Disk Assembly 6 Moog or Marotta $500 $3,000

Pressure Transducer 67 Druck $750 $50,250

Thermocouple 25 Medtherm $200 $5,000

Cavitating Venturi 13 Flow Systems $500 $6,500

0.5 x 0.083 SS Tube 100 m TBD $38/10 m $380

0.25 x 0.039 SS Tube 50 m TBD $18/10 m $90

Docking Connector 8 TBD $20,000 $160,000

Total $4,583,770

The material cost for the propulsion hardware is estimated at slightly over $4.5 million. That does not include assembly or testing, as they are accounted for in the assembly and testing estimates in the cost section.


Final Report 232

Figure A.10.2a – Odyssey Plumbing and Instrumentation (P&ID) Diagram


Final Report 233

Figure A.10.2b – P&ID Legend

A.10.5 – Off-Nominal Maneuver Verification The generous fuel reserves for the altitude maintenance and attitude control systems allow for the capability to perform several types of off-nominal maneuvers that the propulsion system was not necessarily designed for. These include orbital insertion correction and plane change maneuvers. The orbits group has calculated that a maximum of 18.7 m/s of Δv will be required for correction of the Odyssey’s orbit following orbital insertion. Using equation 1 the propellant requirement and burn time for this maneuver was calculated for both the altitude maintenance and attitude control propulsion systems. The 45 degree cant of the ACS thrusters was accounted for in this


Final Report 234

calculation. As seen in table A.10.2 the arcjet can complete the maneuver in 12 days using 42.76 kg of propellant. Six attitude control thrusters can complete the maneuver in 921 seconds using 115.20 kg of propellant.

Table A.10.2 Maximum Orbital Insertion Correction Capability Prop req (kg) Burn Time (s)

Arcjet 42.76 1036257.98 6 ACS Thrusters 115.20 920.81

Plane changes should never be required for the Odyssey as its Sun synchronous polar orbit is ideal for complete earth imaging coverage. However, plane changing is possible using both the AMS and ACS thrusters. The orbits group determined that 135 m/s of Δv is required for a one degree plane change. Using equation 1 the plane change capability of the Odyssey was computed for each propulsion system. Again the 45o cant of the ACS thrusters was accounted for. The results are shown in table A.10.3.

Table A.10.3 – Odyssey Plane Change Capability Arc Jet 6 ACS Thrusters Plane Change (degrees) Propellant Mass (kg)Burn Time (days) Propellant Mass (kg) Burn Time (hours)

0.10 30.89 8.66 83.26 0.18 0.25 77.03 21.61 207.19 0.46 0.50 153.42 43.04 411.16 0.91 1.00 304.35 85.38 809.66 1.80 2.00 598.90 168.00 1570.23 3.49 3.00 883.95 247.96 2284.69 5.07 4.00 1159.81 325.34 2955.82 6.56

Possible using maneuvering fuel reserve Possible using maneuvering and deorbit fuel reserves Possible using all onboard propellant Not Possible

A.10.6 – Auxiliary Propulsion Systems The detailed analysis of the auxiliary propulsion systems is below. The MPS was completed at nearly the same level of fidelity as the rest of the Odyssey systems; however the SMPS and SPS were completed only to an initial conceptual stage. This was due to the need for a higher fidelity analysis in the early stages for the MPS as it will be the primary workhorse for Odyssey. A.10.6.1 – Module Propulsion System For the MPS a structural analysis was done to ensure that the MPS would survive a 7g launch loading. This total g load would be the maximum load that the module would see throughout its lifetime. The MPS is designed to hang from the top of the launch vehicle which leads to all beams being in tension.


Final Report 235

An analysis was performed for both tension and compression loads in the event there are compressive loads due to launch vibrations. For tensile loading equation A.10.6 was used:

xy

FA

σ = (A.10.6)

where Fx is the force along the member, A is the cross sectional area, and σy is the yield stress. In this case titanium was chosen to help eliminate transfer of heat from the MPS to the module, and as such the yield stress is 505.4 MPa. Using a downward load of P = 3000 kg as shown in Figure X.1 and simple geometry the load along the axis, Fx = 34595 N. This leads via, Equation X.1, to a cross sectional area of A = 68 x10-6 m2. For the compressive load case a buckling analysis was done using Equation A.10.7.

2

2

c EFLπ

ρ

=⎛ ⎞⎜ ⎟⎝ ⎠

(A.10.7)

where, F is the axial load, E is young’s modulus, L is the length of the bar, and c is a joint geometry term (for two fixed ends c = 4). For this problem ρ is defined as:

IA

ρ = (A.10.8)

where, A is the cross sectional area, and I is the second moment of inertia for a square beam:

2112

I A= (A.10.9)

The Titanium used has a young’s modulus of 113764 MPa, which when used in Equation A.10.7 results in the cross sectional area A = 1.6 x10-5 m2. This leads to a maximum beam width of 0.004 m.


Final Report 236

Figure A.10.3: MPS Structural Diagram

A safety factor of two was applied to the bar per team requirement which would bring a total bar width to 0.008 m. To reduce cost and time for the “rapid turn-around” vehicle the cost of purchasing specialty machined parts the size was increased to a standard 0.01 m. To approximate the power consumption of the system the power requirements of each component was added up (Table A.10.5) for a total 62W. To operate for the required eight hours on batteries eight 88 Ah NiH batteries will be used. By approximating the mass of the batteries at 4W/kg as discussed in section 11 and noting the required 62W of power a total mass of 15.5 kg is achieved. The solar panel is composed of SunPower 220 cells covering the top of the MPS system. The total area covered by cells is 0.24m2. If exposed to direct sunlight a peak power production of 52.8 W/hr will be reached recharging the entire battery in just over an hour.[ABCD][A.10.7]

The MPS uses a Helium cold gas thruster to capitalize on the extremely low freezing temperature of Helium, eliminating the need for heaters on the fuel tank. To reduce mass a single Helium tank of titanium construction was chosen. Table A.10.4 details the calculations that were performed for the tank and thruster sizing.


Final Report 237

Table A.10.4: Tank and thruster sizing Constants

R 8.3144 J/mol-K Mwt 4.0026 g/mol g0 8 m/s^2 (approx at altitude)

Definitions Temp 5 K (estimated) Pc 50 MPa (desired) Pe 0 MPa (space) Gamma 1.14 (NIST) DV req 210 m/s (to deorbit) DV req 10 m/s (to dock) mdot 0.1 kg/s time 2700 s (burn time) Density Ti 4507 kg/m^3

Propellant Calculations Speed of Sd. 989.64 m/s (NIST) Cstar 1455.96 Isp 467.56 s^-1 Force 374.05 N mprop 250.00 kg Volume 0.05 m^3

Tank Calculations Max T 100.00 K Max P 1.00E+09 Pa Radius 0.23 m Area 0.67 m^2 Thickness 0.23 m Mass 695.21 kg Burst Pres 1.00E+09 Pa (Tank burst Pres) Ftu 5.05E+08 Pa (Yield Stress)

The Helium tanks were sized with a burst pressure of 1GPa to allow for a max full tank temperature of 100 K. This allows for reasonable cryogenic cooling within the launch vehicle and no thermal requirements once the MPS is in space. By not requiring heated tanks less power is needed for the overall system thus reducing complexity, and increasing amount of time that the system can run on batteries alone. As shown in Table A.10.5, additional hardware masses were estimated based on information from other systems within this report.


Final Report 238

Table A.10.5: Component Hardware Mass and Power Estimates Additional Hardware

# Mass (kg) Ttl Mass Power (W) Ttl. Power Valves 12 1 12 1 12 Computer 1 2.5 2.5 10 10 Batteries 1 15.5 15.5 Antenna 3 5 15 10 30 Solar Cells 25 0.05 1.25 Other 100 10

(For references on mass and power see each respective section within this report.) The overall mass of the system with all of the components added together, adds to a total of 1108 kg as shown in Table A.10.6. These masses are proportionally near the SMAD averages for small satellites, lending credibility to the results.[A.10.2]

Table A.10.6: Total System Mass

Total System Mass Propellant Mass 250.00 kg

Tank Mass 695.21 kgStructural Mass 67.03 kg

Additional Hardware Mass 95.75 kg 1107.99 kg

The MPS transmits data from each of three antennas at frequencies of 2.14 MHz, 2.16 MHz and 2.18 MHz. The computer system onboard Odyssey will receive and distinguish each of the frequencies to determine the orientation of the MPS with module. Odyssey will the direct the MPS to maneuver towards the satellite until it is within a distance of 3 km. At 3 km Odyssey’s LADAR system (see Section 7) will take over in determining the orientation of the MPS and will direct the MPS into the docking port. Once docked, the module will be activated and the MPS will be disconnected to maneuver to another module. The MPS will attach to an old or broken module and proceed to undock and de-orbit it. A de-orbit burn will require approximately 210 m/s of ΔV to pass into the atmosphere. The MPS can easily meet this ΔV with an estimated 175 kg of fuel remaining. Overall the primary components of the system were estimated based on information available from internet and professional sources as shown in Table A.10.7. An additional $5000 was added additional miscellaneous hardware.


Final Report 239

Table A.10.7: Component Cost Estimation Cost Analysis

Item Quantity Unit Cost/Unit Total Cost Solar Cells[A.10.8] 96 cells $75.00 $7,200.00 Batteries[A.10.9] 15 kg $100.00 $1,500.00 Computer[A.10.10] 1 each $10,000.00 $14,000.00 Valves 12 each $500.00 $6,000.00 Antenna 3 each $1,000.00 $3,000.00 Structure[A.10.11] 67.0 kg $463.00 $31,034.01 Tank 1 each $80,000.00 $80,000.00 Fuel 270 kg $15.00 $4,050.00 Other $5,000 Total $151,784.01

A.10.6.2 – Super-module Propulsion System For the super-module propulsion system the plan is to scale up the MPS to accommodate the SVT and super module. The additional mass of the super-module requires an additional four bars to accommodate the tensile and compressive loads. The total load applied is approximately 6000 kg and the same equations can be followed as in part A.10.6.4 to determine the applied. To account for the significantly larger mass of the super-module and the low efficiency of helium a hydrazine decomposition thruster system will be used. As a result of this there will need to be additional thermal heaters to keep the hydrazine from freezing. Also there will be additional power requirements for the added heaters, valves, and additional hardware.


Final Report 240

A.10.7 – References [A.10.1] Hrbud, Ivana. A&AE 590E Class Notes. Purdue, 2006. [A.10.2] Larson, Wiley, and James Wertz. Space Mission Analysis and Design. Microcosm,

1999. [A.10.3] Sutton, George P., and Oscar Biblarz. Rocket Propulsion Elements. 7th ed. New York:

John Wiley & Sons, 2001. [A.10.4] Humble, Ronald, Gary Henry, and Wiley Larson. Space Propulsion Analysis and

Design. McGraw-Hill, 1995. [A.10.5] Potts, C.L., Raofi, B., and Kangas, J.A., Mars Exploration Rover Propulsive Maneuver

Design, AIAA 2004-4985. [A.10.6] Brown, Charles D. AIAA Element of Spacecraft Design: Home Study Correspondence

Course. 2001 [A.10.7] "SPR-220 High Efficieny PV Module." SunPower. 18 Nov. 2006

<http://www.sunpowercorp.com/home.html>. [A.10.8] "SunPower." Advertisement. 20 Nov. 2006 <sunpowercorp.com>. [A.10.9] Zhang, Hannah. "Purdue Solar Racing Battery Quote." E-mail to Matt Hewlett.

15 June 2006. [A.10.10] "Rugged Computer Systems." AiTech. 2006. 7 Nov. 2006 <http://www.rugged.com>. [A.10.11] "Titanium." McMaster-Carr. 25 Nov. 2006. 25 Nov. 2006

<http://www.mcmaster.com>.


Final Report 241

A.11. – Power System Appendix A.11.1 - Solar Panel Equations: The required power output needed by the solar array, not including degradation, is given in equation 1 below.

1−−−

+=a

d

dlbba

nnsa X

PTXX

TPP (A.11.1)

Working with excel, a table was made to calculate this Psa and is given as Table A.11.1 below.

Table A.11.1-Solar array power calculations

Total power req from other subsystems 3,311.70Watts 30% power loss over 10 years 4305.21Power needed to generate in yr one Solar Array Calculations Average power output from solar arrays over daylight period Psa 7061.455556 spacecraft daylight,time in sunlight hours Td 16.8hours spacecraft night, hours Tn 7.2hours average power consumed in night Pn 4305.21 average power consumed in day Pd 4305.21 power transfer efficiency, solar array to daytime loads Xa-1 0.9 power transfer efficiency, array to batt., in batt. Charge efficiency. Xa-b 0.9 power transfer efficiency, batt to nightitme loads Xb-1 0.97

This table gives some of the tabulated results for the power output from the solar array. Furthermore, it shows the increase in power needed for the end of life operation which is seen in the 30% power loss over 10 years cell. A.11.2 - Solar area calculations Solar area was calculated via certain efficiencies and assumptions given in Table A.11.2 below.

Table A.11.2-Solar array assumptions Power from on cell under operational cond. Pc 0.0490 watts / cell

Power delivered from once cell lab cond Pl 0.0978 Watts/ cellpower loss due to UV discoloration nuv 0.9800 power loss due to thermal cycling ncy 0.9900 power loss due to cell mismatch nm 0.9750

power loss due to resistance in cell interconnect nl 0.9800 power loss due to containmation ncon 0.9900

loss due to shadowing ns 1.0000 power loss due to radiation damage nrad 0.9800

power adj for temp nt 0.9900 Adj. for solar intensity Hi 0.5650

array pointing loss factor Lp 0.9960


Final Report 242

With these efficiencies the actual area was calculated via equation A.11.2.

c

csa D

NA = (A.11.2)

Nc is the number of cells required to provide necessary power requirement at the end of life and is given by equation A.11.3 below

c

sac P

PN = (A.11.3)

Dc is the cell density and is given by equation A.11.4 below

88.10000

cellc A

D = (A.11.4)

The packing factor for our solar array is the .88 in the equation above and Acell is the individual area of one cell in the solar array which is at 4 cm2. A.11.3 - Battery calculations The battery energy was calculated using equation A.11.5 listed below

lb

nnb X

TPE

−

= (A.11.5)

Pn, Tn, and Xb-l are the same as they were for the solar array calculations After picking a depth of discharge for the battery, as they are never allowed to be completely drained, the actual energy output for the battery was obtained from the relation given by equation A.11.6

DODXTP

Elb

nnBCap

−

= (A.11.6)

Where DOD is the depth of discharge The amp hour capacity was figured next and from this the number of batteries could be conceived based on the fact that 80 amp-hour capacity batteries would be used

d

BCap

VE

C = (A.11.7)

Where Vd is the battery voltage discharge which is 28 volts


Final Report 243

Mass of the battery system was found using equation 8 below

ergySpecificEnE

Mass b= (A.11.8

Table A.11.3 below gives the numerical results for the batteries

Table A.11.3-Battery results Battery Calculations Battery Discharge rate Vd 28volts Depth of Discharge DOD 0.3 Amp hour capacity C 272.9944158amp-hour Number of batteries 3.412430198 Eb 2224.3585Watt-hour weight of batteries 92.68160417


Final Report 244

A.12 – Thermal Control Table 12.1 lists the important values that correspond to the orbit. These values are the values that affect the thermal environment model. The first four values are given from the orbits group. The fourth value is the difference between the Right Ascension of the Ascending Node and the Right Ascension of the Sun. For a sun-synchronous orbit, the difference is constant. The Orbit Beta Angle is calculated from equation 12.1 and 12.2. The minimum beta angle is calculated in equation 12.3. It is the beta angle at which the satellite does not enter Earth’s Shadow and has a zero eclipse fraction. The eclipse fraction is calculated in equation 12.4. In Graph A.12.1 the Orbit Beta Angle and eclipse fraction is shown for one year

Table A.12. 1 Orbital Parameters Earth’s Radius (km) R 6378

Altitude (km) h 785 Inclination (deg) RI 98.54

RAAN-RAS (deg) Ω-ΩS 146 Orbit Beta Angle range (deg) β 27 to 35

Minimum beta angle with no eclipse (deg) β* 62.8 Eclipse fraction 30%

Period (Minutes) 100

)cossin)sin(sin(cossin 1 RIRI SSS δδβ +Ω−Ω= − (A.12.1)

))10(365

360cos(*45.23 daysNdaysS +−=δ (A.12.2)

Equation 12.2 is for the Sun’s Declination where N is day of the year (Jan 1 = 1, July 4th=185)

⎥⎦⎤

⎢⎣⎡

+= −

hRR1* sinβ (A.12.3)

*

*2/12

1

0

cos)()2(cos

1801

ββ

βββ

≥=

<⎥⎦

⎤⎢⎣

⎡++

= −

if

ifhR

Rhhf E (A.12.4)


Final Report 245

Graph A.12.1: Orbit Beta Angle for One Year

0.000

5.000

10.000

15.000

20.000

25.000

30.000

35.000

40.000

0 20 40 60 80 100

120

140

160

180

200

220

240

260

280

300

320

340

360

Days

Orb

it B

eta

Ang

le (D

eg)

Eclip

se F

ract

ion

(% o

f orb

it))

Orbit Beta Angle

Eclipse Fraction

With the values from Table 12.1 albedo and Earth IR values can be chosen. The albedo and Earth IR are chosen from Figures A.12.2 and A.12.3. I used an average beta angle of 30 and orbit angles from 0 to 40 from Figure A.12.1 for ρ values of 180º, 135º and 90º. For the albedo values, I picked 5 values at each of the three angles and averaged them. Table A.12.2 has these values

Figure A.12. 1

Table A.12. 2 Albedo values 90º 135º 180º

θ = 40 75 180 260 θ = 30 80 215 300 θ = 20 85 225 315 θ = 10 85 230 330 θ = 0 90 235 350

Average 83 217 311


Final Report 246

Figure A.12. 2 Albedo values

Earth IR values that were selected from Figure A.12.3 are 166 W/m2 for 180º surfaces, 120 W/m2

for 135º surfaces and 50 W/m2 for 90º surfaces. Table A.12.3 lists the elements of the insulation and the summary of the weight and thickness.

Table A.12. 3 Insulation Layers

Name # of

LayersThickness mm (mil)

Weight (gm/cm2)

Outer Layer


1 0.254 (10) 0.0036

Middle Layers

Aluminized Mylar 10 0.00635 (.25) 0.00093

Inner Layers


1 0.0254 (1) 0.005

Spacers Dacron 11 0.1651 (6.5) 0.00065 Total 12 2.159 (85) 0.02483


Final Report 247

Figure A.12. 3 Earth IR Values

The next set of calculations is used to size the heaters for the propulsion segment. There are three tanks of hydrazine and three tanks of helium. There are also 90 meters of propellant lines, 12 thrusters and one arcjet. The steps are as follows 1) Obtain tank volume and calculate the tank radius including the thickness 2) Compute the surface area 3) Compute the power lost for a temperature of 283 K and emissivity of 0.01 4) Size the heaters assuming that there would be six heaters with twice as much power as the

tanks emit. The equations used here for emissivity involves the Stefan-Boltzmann Law . Radiative heat transfer is dependant upon the emissivity times the Stefan Boltzmann constant (σ) times the absolute temperature to the fourth power. The results for the hydrazine and helium tanks are shown in table A.12.5.

4TQ εσ= (A.12.5)

propellantofmassCAreaQT

P **

=•

(A.12.6)


Final Report 248

Table A.12. 4 Crucial Propulsion Values Temp of tanks (K) 283

Hydrazine mass (kg) 1019.67 Hydrazine g/mol 32

Tank MLI emissivity 0.01 Stephan-Boltzmann Constant (W/m2 k4) 0.000000056

Cp of Hydrazine (J/kg K) 3297.166781 Cp of Titanium (J/Kg K) 522

Volume of each hydrazine tank (in3) 28175.18894 Thickness of hydrazine tank (in) 0.094069934

Hydrazine tank mass (kg) 36.20679298 Cp of Helium (J/kg K) 5196.5

Volume of He tank (ft3) 6.513473872 Thickness of Helium Tank (in) 0.692809815

Helium tank mass (kg) 128.3093658 Helium mass (kg) 243.4318079

Conductivity of Titanium alloy (W/m K) 7.8

Table A.12. 5 Heater Size Calculations Hydrazine Helium

Volume of one tank (in3) 28175.19 11255.28 Radius (in) 18.97 14.60 Area (in2) 4522.51 2676.88

Heat out (W) 10.48 6.20

Temp Change (K/s) -

0.000003117 -0.000004903 Time for 2° temp drop (days) -7.43 -4.72

Time for heaters to reach 285 K (days) 3.99 2.48 Time for full tanks to drop from 283 to 275 (Days) 29.71 18.89

Heater Size (W) 5 3


Final Report 249

Table A.12. 6 Environmental Loads on the Modules

Satellite in the Sun Component

Mass Angle to the Sun

Angle to the Earth

Absorbance Factor

Emittance Factor

Solar Load

Earth IR Albedo

Super Module Front 68.500 30 180 0.14 0.40 166.2 0.0 0.0 Super Module Side 68.500 60 90 0.14 0.40 96.0 20.0 0.0

5 W Module 12 90 45 0.30 0.31 0.0 37.2 65.1 20 W Module 135 90 45 0.30 0.31 0.0 37.2 65.1 40 W Module 12 90 45 0.30 0.31 0.0 37.2 65.1

90 Watt Module 245 90 0 0.14 0.40 0.0 66.4 43.5 100 W Module 34 90 45 0.14 0.40 0.0 48.0 30.4 200 W Module 50 90 45 0.07 0.79 0.0 94.8 15.2

Satellite in the shadow Angle to the Sun

Angle to the Earth

Absorbance Factor

Emittance Factor

Solar Load

Earth IR Albedo

Super Module Front 68.500 90 90 0.14 0.40 0.0 0.0 0.0 Super Module Side 68.500 90 90 0.14 0.40 0.0 20.0 0.0

5 W Module 12 90 45 0.30 0.31 0.0 37.2 0.0 20 W Module 135 90 45 0.30 0.31 0.0 37.2 0.0 40 W Module 12 90 45 0.30 0.31 0.0 37.2 0.0

90 Watt Module 245 90 0 0.14 0.40 0.0 66.4 0.0 100 W Module 34 90 45 0.14 0.40 0.0 48.0 0.0 200 W Module 50 90 45 0.07 0.79 0.0 94.8 0.0

Satellite in the Sun Total

absorbance Emittance Temp Electronics

Power Heater power

Net Power Gain

Super Module Front 191.2 139.4 280.024 25.0 0.0 -51.80 Super Module Side 171.0 141.1 280.842 25.0 30.0 -29.9

5 W Module 137.3 113.6 283.551 5.0 30.0 -23.7 20 W Module 122.3 111.6 282.300 20.0 0.0 -10.7 40 W Module 142.3 129.3 292.877 40.0 0.0 -13.0

90 Watt Module 199.9 164.0 291.604 90.0 0.0 -36.0 100 W Module 178.4 157.0 288.467 100.0 0.0 -21.4 200 W Module 310.0 309.9 288.424 200.0 0.0 -0.1

Satellite in the shadow Total

absorbance Emittance Temp Electronics

Power Net Power Loss

Super Module Front 25.0 145.9 283.2 25.0 0.0 139.4 Super Module Side 75.0 144.8 282.7 25.0 30.0 141.1

5 W Module 72.2 127.5 291.8 5.0 30.0 113.6 20 W Module 87.2 112.1 282.6 20.0 30.0 111.6 40 W Module 107.2 137.5 297.4 40.0 30.0 129.3

90 Watt Module 81.4 165.4 292.2 0.0 15.0 164.0 100 W Module 113.0 162.8 291.1 50.0 15.0 157.0 200 W Module 309.8 309.9 288.4 200.0 15.0 309.9


Final Report 250

Table A.12.7 Total Thermal Mass and Power

Device Weight(kg)

Power(W)

Back upQuantity

PrimaryQuantity

Surface Area (m2)

Total Mass (kg)

Total Energy (W)

Propulsion tank heaters 0.088 5 18 18 3.17 90Propellant line heaters 0.088 1 45 45 7.93 45Thruster heaters 0.088 1 13 13 1.15 13Propellant thermostats 0.030 0.1 152 4.56 15.2Propulsion tank insulation 0.248 3 2.92 2.17 0Helium tank heaters 0.088 3 18 18 1.59 54Helium tank thermostats 0.030 0.1 36 1.08 3.6Propulsion segment outer surface 0.009 1 30.63 0.28 0Module thermostats 0.030 0.1 30 0.9 3Module heaters (large) 0.088 30 9 9 1.59 270Module heaters (small) 0.088 15 6 6 1.06 90Module inner insulation 0.248 14 3.79 13.19 0Core insulation 0.248 1 38.03 9.44 0Super module outer surface 0.020 1 80.2 1.60 0Module outer surface paint 0.240 8 1 1.92 0Module outer surface radiator 0.020 4 1 0.08 0Quartz mirror 1 2 1 2 0Final Weight (kg) 53.70 583.8


Final Report 251

A.14 – Cost and Schedule

Table A.14.1 – Cost Analysis Calculation Table

The CER values, which are the cost of the components in thousands of dollars, are computed in Table A.14.1. The following equations were used from SMAD. 1.1, 1.2: 44263(X0.562) = CER 3: .0052(X1) = CER 1.4: 20+230(X0.59) = CER 4: 0.11(X1) = CER 1.5: 179*X = CER 5: 0.36(X1) = CER 2: 16253+110(X0.77) = CER 2.1: 86(X0.65) = CER 2.2: 93+164(X0.93) = CER 2.3: 1244(X0.39) = CER 2.4: 183(X0.29) = CER


Final Report 252

the odyssey project team members - purdue university · aae450 - senior spacecraft design – fall...

Documents