AOE 4065

Space Design

C U B I K

CubeSat Universal Bus Integrated Kit Research and Design

22 November 2002

Submitted by the NeoCubists:

David Carton Fadi Mantash Julien Pierru

Ryan Reisman Ankit Singhal

Robert Thompson Bryan Tisinger

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Table of Contents

Table of Contents………………………………..………………………………………...ii List of Figures….………………………………..………………………………………..iv List of Tables………..…………………………..………………………………………...v List of Abbreviations…..………………………..………………………………………..vi List of Symbols………………………………..…………………………………………vii

Chapter 1: Introduction and Problem Definition 1.1 History and Background……………………………………………………….….1 1.1.1 Satellite History………………………………………………………………. .1 1.1.2 Pico Satellite History…………………………………………………………. .1 1.2 Problem Definition…………………………………………………………….…..2 1.2.1 Introduction……………………………………………………………………..2 1.2.2 Scope……………………………………………………………………………4 1.2.3 Disciplines…………………………………………………………….………...5 1.2.4 Societal Sectors and Actors Involved…………………………………………..5 1.2.5 Needs, Alterables, and Constraints……………………………………………..7 1.2.6 Relevant Elements……………………………………………………………...9 1.3 Summary…………………………………………………………………………..9 Chapter 2: Value System Design 2.1 Project Objectives………………………………………………………………..11 2.1.1 Performance Objectives……………………………………………………….12 2.1.2 Cost Objectives………………………………………………………………..14 2.1.1 Simplicity Objectives………………………………………………………….16 2.2 Decision Hierarchy………………………………………………………………18 2.3 Summary……………………………………………………………………........19 Chapter 3: System Synthesis 3.1 Structure……………………….............................................................................20 3.1.1 Material………………………..........................................................................20 3.1.2 Structural Elements………………………........................................................21 3.1.3 Internal Structure………………………...........................................................22 3.1.4 Control Elements, Bus Mechanisms, and Deployables…………………….....23 3.1.5 Thermal Effects………………………..............................................................34 3.2 Flight Computer……………………….................................................................34 3.2.1 The Control Unit………………………............................................................25 3.2.2 The Data Storage………………………............................................................25 3.3 Flight Computer Software………………………..................................................26 3.3.1 The Real Time Operating System………………………..................................26 3.3.2 The Micro Controller Case………………………............................................27 3.3.3 The Microprocessor Case………………………..............................................27 3.3.4 The Drivers for Subsystem Devices………………………..............................27 3.4 Communications………………………................................................................28 3.4.1 Transponders………………………..................................................................28 3.4.2 GPS Receivers………………………...............................................................30

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3.4.3 Antennas………………………........................................................................31 3.5 Attitude Determination and Control System……………………….....................33 3.5.1 Orbit Attitude Stability………………………..................................................33 3.5.2 Causes of Attitude Instability………………………........................................34 3.5.3 Attitude Determination………………………..................................................36 3.5.4 Attitude Control……………………….............................................................37 3.5.5 ADCS Best Options……………………….......................................................40 3.5.6 ADCS Summary………………………............................................................41 3.6 Power……………………….................................................................................41 3.6.1 Power Generation…………………..................................................................41 3.6.2 Power Storage...............................………………………................................42 3.6.3 Power Regulation and Distribution...................................................................43 3.7 Documentation………………………..................................................................44 3.7.1 Documentation for TransOrbital, Inc………………………............................45 3.7.2 Documentation for the Users……………………….........................................45 3.8 Summary………………………............................................................................46 Chapter 4: System Analysis 4.1 Structure……………………….............................................................................47 4.1.1 Material………………………..........................................................................48 4.1.2 Structural Elements………………………........................................................49 4.1.3 Internal Structure………………………...........................................................49 4.2 Flight Computer……………………….................................................................50 4.2.1 The Storage Unit………………………............................................................50 4.2.2 The Processing Unit………………...................................................................51 4.2.3 The Control Unit................................................................................................52 4.3 Flight Computer Software………………………..................................................52 4.3.1 The Operating System………………………...................................................53 4.3.2 The Programming Language…………………….............................................54 4.4 Communications………………………................................................................55 4.4.1 Transponders………………………..................................................................55 4.4.2 Antennas………………………........................................................................56 4.4.3 GPS Receiver.....................................................................................................57 4.5 Power……………………….................................................................................58 4.5.1 Power Generation…………………..................................................................58 4.5.2 Power Storage...............................……………………….................................59 4.6 Attitude Determination and Control System……………………….....................60 4.6.1 Attitude Determination………………………..................................................60 4.6.2 Attitude Control……………………….............................................................61 4.7 Documentation…………………...........................................................................62 4.7.1 Documentation for TransOrbital, Inc……………………….............................63 4.7.2 Documentation for the Users……………………….........................................63 4.8 Summary………………………............................................................................64

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List of Figures

Figure 2.1: The VSD Objectives Form a Hierarchy…………………………………….11 Figure 2.2: The Sub-level Objectives for Maximizing Performance…………………....13 Figure 2.3: The Sub-level Objectives for Minimizing Cost.…………………………….15 Figure 2.4: The Sub-level Objectives for Maximizing Simplicity….…………………...17 Figure 3.1: Possible Rail Post Configurations.………………………………………......22 Figure 3.2: PC/104 Form Factor Dimensions….………………………………………...23 Figure 3.3: The Three Forms of Attitude Stability………………………………………34 Figure 4.1: A Generalized Representation of the Flight Computer.......…………………50 Figure 4.2: The Communication Subsystem Components and Interactions..……………55 Figure 4.3: The Performance Characteristics of the Battery Types...……………………59

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List of Tables

Table 1.1: Needs, Alterables, and Constraints ……………………………………………8 Table 2.1: Project Objectives and Their Relevant Subsystems………………………….12 Table 2.2: Sub-level Objectives for Minimizing Cost……………………………….…..14 Table 2.3: Decision Hierarchy Weights………………………………………………….18 Table 3.1: Attitude Determination Sensors....……………………………………………37 Table 3.2: Attitude Control Options for CubeSat CUBIK System...…………………….38 Table 4.1: Material Properties of the Alloys Considered for the Structural Elements......48 Table 4.2: Characteristics of Memory Devices................................……………………..51 Table 4.3: Intel Processor Operating Characteristics.......................……………………..52 Table 4.4: Operating System Alternatives and Specifications..........……………….........53 Table 4.5: Transponder Characteristics...........................................……………………..55 Table 4.6: Antenna Characteristics..................................................……………………..56 Table 4.7: Communications Summary............................................……………………..57 Table 4.8: Relative Weights for ADCS Determination of the Three Models.....………..62 Table 4.9: Documentation Choice for Each Design Model............……………………..64 Table 4.10: Summary of the Three CUBIK Design Models.............................................65

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List of Abbreviations

ADCS Attitude determination and control system

CUBIK CubeSat Universal Bus Integrated Kit

CPU Computer Processing Unit

CU Control Unit

FCS Flight Computer Software

FM Frequency Modulation

GPS Global Positioning Sensors

GUI Graphical User Interface

I/O Input/Output

LEO Low Earth Orbit

MOE Measure of Effectiveness

NiCd Nickel Cadmium

OS Operating System

OSSS One Stop Satellite Solutions

PC104 PC Form Factor 104

POM Polyoxymethylene

PROM Programmable Read Only Memory

RFP Request for Proposal

RTOS Real Time Operating System

USB Universal Serial Bus

VSD Value System Design

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List of Symbols °C Celsius, degrees

cm Centimeters

a Coefficient of Thermal Expansion

$ Dollars (US)

“ Inch

kg Kilogram

m Meter

µm Micrometers

mm Millimeters

Q Proportional

∴ Therefore

l Wavelength

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Chapter 1: Introduction and Problem Definition

1.1 History and Background

The impact of satellites on everyday life can be seen everywhere. From allowing

effective global communications to aiding in weather forecasting, satellites enrich society

by providing access to Earth orbit. What advances might be developed if satellite design

and launching were simplified? Recent developments in satellite technology now have

the potential to answer the question. A new generation of small-scale satellites offers the

promise of simple, cost effective, ready-made access to space.

1.1.1 Satellite History

Since the launch of Sputnik on October 4, 1957, the domain of space has been

dominated by governments and big business. Because of the high cost of putting a

satellite in Low Earth Orbit (LEO), smaller businesses, universities or other

educational/research facilities must really on grant money to put a full scale satellite in

orbit. Costs can range into the millions of dollars (US) for a single launch. The cost is

often beyond the resources of grant contributors1. However, there is a solution. The cost

of the addition of a small secondary payload on primary payload launch is considerably

less2. “Piggy-backing” as a secondary payload forms the basis for creating the

picosatellite industry.

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1.1.2 Picosatellite History

The need of greater public access to space science brought the original design of

the picosatellite to the forefront a few years ago. Research done at California Polytechnic

State University (Cal Poly) and Stanford University among others created a simple usable

design for a picosatellite that could be launched as a secondary payload on many major

satellite missions3. Companies such as One Stop Satellite Solutions Inc. (OSSS) market

these picosatellites as kits with predefined missions4. Customers have a limited range of

choices for missions and may not be able to afford to redesign the kits to fit a desired

mission. This need for versatility of missions for the customer is the focus of this project.

1.2 Problem Definition

The project begins with defining the problem. Explaining the motivation for a

design solution is then presented. An overview of the project is given by discussing the

factors which should be addressed. An emphasis is placed on pairing the factors involved

in the project with the people who interact with the factors.

1.2.1 Introduction

Currently, the resources required to design and build a satellite are beyond those

available to many who wish to study and explore space. Providing a low-cost and

functionally complete satellite product would revolutionize access to space. TransOrbital,

Inc., in conjunction with Virginia Tech, has proposed developing a satellite kit to address

the potential market for affordable space access. Taking advantage of recent

developments in picosatellite technology, the CubeSat Universal Bus Integrated Kit

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(CUBIK) senior design group intends to design and build a functional prototype based on

the CubeSat project architecture.

CubeSat is a standardized micro-spacecraft architecture developed by Professor

Robert Twiggs of Stanford University, in association with Cal Poly. The CubeSat is a

picosatellite measuring approximately 100 millimeter per side5. The entire satellite,

including payload, has a total mass of less than 1 kilogram. The small size and low mass

allow many launch vehicles to carry CubeSats as secondary payloads. Launching

CubeSat picosatellites as secondary payloads lowers cost. In addition, the construction

cost for a picosatellite employing the CubeSat architecture is within reach of many

universities, schools, amateurs, and small business design teams.

Although the CubeSat architecture provides an effective low-cost design for a

picosatellite, the manufacturing resources and technical knowledge required for

construction are still prohibitive. The aim of the CUBIK project team is to design a

functional picosatellite kit which will be marketed to customers lacking the resources to

design and build a satellite. For example, a public school might purchase a CUBIK kit as

a means for teaching children about space. CUBIK kits equipped with a GPS receiver

would allow the students to track the satellite’s position in orbit and know when to listen

for the satellites transmissions.

The CUBIK satellite kit is an affordable test platform for miniature space-rated

components such as gyros, momentum wheels, and position sensors. Currently, there are

few manufactures of such systems. Another scenario is an aerospace firm purchasing

CUBIK kits to test their space-rated products and new technology.

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1.2.2 Scope

The project is the development of an affordable and reliable universal picosatellite

bus system which meets the specifications of TransOrbital, Inc. TransOrbital has defined

the scope of this project as listed below:

• Designing an affordable commercial kit

• Providing documentation to produce the kit including specifications and source

code

• Determining the appropriate communications frequencies

• Deciding power usage for the payload and bus

• Designing the computer and control systems

• Defining the maximum altitude communications can be maintained

Scope factors which have not been set by TransOrbital include:

• Launch vehicle

• Ground control

• Deployment system

• Payload function

• Manufacturing subcomponents

The scope outlined above determines the project goals.

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1.2.3 Disciplines

A variety of disciplines are involved in the completion of a major project such as

a CubeSat kit. An assortment of engineering fields including aerospace, mechanical, and

electrical must work cooperatively throughout the duration of the project to develop the

CubeSat bus. Aerospace engineers will be responsible for the attitude and position

sensors. Verifying the suitability of components for use in space is done by systems

engineers. Mechanical engineers will test the spacecraft’s mechanical structure. Finally,

the power subsystems and the communication subsystems will be designed and tested by

electrical engineers.

In addition to these engineering disciplines, scientists will participate in this

project. Computer scientists will be responsible for the flight computer, its software

interface, and programming. Materials specialists may be consulted concerning materials

selection.

1.2.4 Societal Sectors and Actors Involved

The CUBIK project will involve many societal sectors. CUBIK is intended to

provide inexpensive access to space for currently disenfranchised sectors of society. A

principal sector is the education community. The CUBIK kit is intended to allow high

schools, colleges, and science clubs to build and launch satellites. Inexpensive kit

satellites could also create more opportunity for aerospace students to operate “hands-on”

with functional space hardware.

The aerospace community is another targeted sector. Small inexpensive satellites

could create more interest in space technology and space research. Expendable

picosatellites could be used to safely test new satellite materials, solar cells or propulsion

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devices without the concern of damaging an expensive satellite bus or jeopardizing a

mission.

The CUBIK project could have an impact on the business community as well.

CUBIK may allow small businesses the chance to launch satellites for commercial

purposes. CUBIK kits might create new satellite business. A picosatellite bus kit could

increase the number of annual commercial satellite orders. Currently, annual commercial

satellite orders only number in the dozens6.

The impact on the environmental community would be minor. CubeSat satellites

are ideally secondary payloads. No increase in rocket launches is anticipated. The most

significant concern is the impact on the space environment. The space environment

would be affected by the increased amount of space debris. Another environmental factor

would be the need for frequency bands for communications. Satellite communications

frequencies are a regulated resource7.

The actors involved in the design project are both individuals and institutions. The

actors have either academic or commercial interest in the development of CubeSat

picosatellite technology. The principal actors are listed below.

The Virginia Tech CUBIK Design Team (The NeoCubists) is responsible for

developing a CubeSat based satellite kit. Advising the CUBIK team is Virginia Tech

professor, Dr. Chris Hall. In addition to offering technical guidance, Dr. Hall acts as

liaison between the CUBIK team and the project sponsor, TransOrbital, Inc.

TransOrbital, Inc. has requested a commercial product design based on the specifications

developed by Cal Poly and Stanford University.

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1.2.5 Needs, Alterables, and Constraints

The design objectives and limitations are defined by the needs, alterables, and

constraints (NACs). The NACs come various sources, including the Request for Proposal

(RFP) provided by TransOrbital, Inc. and the CubeSat specifications developed by Cal

Poly and Stanford University. Other NACs are imposed by the actors listed in Section

1.2.4.

Table 1 lists the needs, alterables, and constraints relevant to CUBIK design. The

goal is to create a satellite bus system kit based on CubeSat architecture. The bus needs to

accommodate a variety of customer payloads and affordable. Payloads will be determined

by purchasers of the bus kit, therefore the bus will need to be stable with a range of

possible performance requirements.

Designs for a universal bus must address power availability. The selection of bus

subcomponents will determine the power available for the payload, as well as power

generation and storage. The bus will need a computer to control electronics. The design

of computer and how the customers access it are project alterables. The design team must

decide the electronics functionality, considering economic factors and customer needs.

There are numerous constraints for the project. Power system constraints include

a stipulation that solar cells must occupy at least five of the six cube sides. An access port

must be included for battery testing and charging. An electrical “kill” switch must be

installed on the structure. There are also structural constraints. The bus interaction with

the launcher and other satellites is specified, including the minimum surface contact.

Aluminum is recommended for the body of the bus, so that the satellite has the same

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thermal expansion characteristics as the launch mechanism developed by Stanford

University.

Table 1.1 Main Objective, Needs, Alterables, and Constraints

Category Element

Main Objective

Design Universal Bus System based on CubeSat

Needs Affordable to companies, colleges and schools

Reliable without backup systems

Ease of manufacture

1 year life span

Alterables Available power and control interfaces

Location of subsystem components

Attitude control system

Subsystem component selection

Control system capacity

Programming language

Additional ports/interfaces available

Retirement capability

Frequencies available for use

Component and material costs

Orbit

Debris and radiation resistance

Constraints Interior volume of 50% or more available for payload

<1 kg mass, including payload

Can not generate debris

Must meet all CubeSat specifications

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1.2.6 Relevant Elements Several subsystems of the CUBIK design interact. The subsystem interactions are

described as relevant elements. The interactions determine the limitations and

requirements of each subsystem. For example, power is a principal concern for CUBIK.

The available power determines the versatility of the payload and the bus. The more

power the bus can provide, the more functional the payload can be. More available power

provides for longer transmission times for the communication system, more computer

operations, and longer payload operation.

Additional elements of the CUBIK design include the spacecraft mechanical

structure, the flight computer, software, communication subsystem and the attitude and

position sensors. The communications interact with the computer and make demands on

the power system. The thermal radiation from the electronics might interact with the

structure as well as the payload.

Cost is a relevant element. The budget determines the type of materials and

systems which can be utilized in the design. Cost impacts prototype construction and the

manufacturing of the finish product. Finally, retail cost is important. The product must be

affordable for the customer base specified in the RFP.

1.3 Summary

Access to space is limited by cost and technical factors. Many academic and small

business groups lack the resources to design, build and launch a satellite. Picosatellite

technology provides a low-cost alternative for groups wishing to launch a satellite. The

CubeSat architecture pioneered at Stanford University has the potential to be adapted for

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commercial markets. A standardized satellite bus kit based on CubeSat would provide an

affordable, relatively non-technical means for institutions to access space.

The goal of the CUBIK project is to design a picosatellite kit design based on the

CubeSat architecture. The design will incorporate both the requirements listed by the

TransOrbital, Inc. request for proposal and the Stanford University CubeSat specification

document. The following chapters continue the problem definition discussion, further

addressing the needs, alterables, constraints, relevant elements interactions, and actors

involved in the design process.

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Chapter 2: Value System Design 2.1 Project Objectives

The Value System Design (VSD) process relates needs and constraints with the

CUBIK design objectives. VSD organizes the CUBIK project objectives into a hierarchy,

as seen in Figure 2.1. The main objective is to design a universal affordable CubeSat bus

kit. The three top-level objectives under the main objective are: maximize performance,

minimize cost, and maximize simplicity. Table 2.1 lists the sub-level objectives for the

project.

The VSD process includes determining measures of effectiveness (MOE) for each

objective. MOEs quantify the design of the sub-level objectives. In addition, the MOEs

are used to evaluate alternative designs in later chapters.

Figure 2.1 The VSD objectives form a hierarchy.

Legend:

Design Universal Affordable CubeSat Bus

Cost

Simplicity

Performance

Maximize

Minimize

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Table 2.1 Project Objectives and Their Relevant Subsystems

2.1.1 Performance Objectives The performance aspect of the objective hierarchy is further divided into

sublevels, which can be viewed in Figure 2.2 below. Measures of effectiveness are a way

to evaluate the effectiveness of the objectives. Performance MOEs consist of various

quantities and units. The performance of the mass and structure of the CubeSat are

characterized by quantities such as mass in kilograms, volume in cubic millimeters, and

vibration test results in hertz.

Objective Description Relevant Subsystems(s)

Maximize Performance Maximize power efficiency Power, payload Minimize power consumption of Bus---(max power to payload)

Power, communication, attitude and control, computer

Minimize mass of bus Structure, power Maximize payload volume Structure Max stability Attitude and control, computer, power, structure Max communication range Power, communication, payload Minimize interference with other satellites Structure, computer, power Maximize versatility for customer experiment

Structure, computer, software, communication, power

Maximize computer efficiency Power, payload Maximize structure Minimize thermal effects Structure, power, communication, computer Minimize radiation effects

Computer, communication

Minimize Cost

Minimize research and development costs

Computer, communications, power, structure, attitude and control

Minimize production costs All Minimize use of public domain codes

Software

Maximize Simplicity

Maximize clarity of documents for user All Maximize simplicity of computer interface to program the kit

Computer, software

Maximize the simplicity of interfaces between bus and payload

All except structure and attitude and control

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Figure 2.2 The sub-level objectives for maximizing performance. Maximizing or minimizing the sub-level objectives characterizes the performance

of CUBIK. Specifically, the performance of the CUBIK power system is measured by

wattage. The power consumption of the bus needs to be minimized whereas the power

efficiency must be maximized. The bus mass must be minimized to have an effective

design. There is a mass constraint of <1 kilogram. The lower the mass of the bus, the

more available mass there is for the payload. Similarly, the there is a constraint that half

Performance

Power Efficiency

Structural

Stability

Communications

Versatility for User

Allowable Payload

Computer

Power Gain (dB)

Output / Input

Stress Strain

Pointing Error (m)

CPU Speed (MHz)

Volume (m3)

Design/ Satisfaction Questioner

Bus Power Use

Power (W)

Mass

Thermal Effects

Potential CubeSat Interference

Mass (kg)

Temp. (°C)

Activation Time (s)

Coefficient of Thermal Expansion

MOE

Minimize

Maximize

Legend

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the internal volume of the bus must be reserved for the payload. The internal volume

available for payload is a MOE for the bus kit performance.

There are other performance objectives, each with its own MOEs. Results from a

stability test will determine the effectiveness of the CUBIK structure. The performance of

the electrical systems has a variety of MOEs. Communication range is measured by

power gain. The computer efficiency is quantified by bus speed and data storage. The

potential to cause interference with other CubeSats is gauged by activation delay time in

seconds. Finally, the sensitivity of the electronics to both thermal and radiation effects

are measured by the internal temperature and shielding efficiency, respectively.

2.1.2 Cost Objectives One of the objectives of CUBIK is to minimize the cost. The sub-level objectives

under the cost objective are presented in Figure 2.3. To minimize the cost, there are three

sub-level objectives: minimize research and development, minimize the production cost,

and use existing software code. The sub-level objectives are summarized in Table 2.2.

Table 2.2 Sub-level Objectives for Minimizing Cost Objective Reasoning Minimize research and development costs

Use of retail components is cheaper than custom made components

Minimize the production costs Components which are rare and hard to manufacture increase production cost

Maximize use of public codes and standards

Avoiding software licensing requirements lowers cost. Using standard I/O codes avoids need to develop new codes.

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Figure 2.3 The sub-level design objectives for minimizing cost.

Minimize research and development cost: To minimize the cost, research and

development cost needs to be minimized. CUBIK relies on built and tested parts rather

than parts designed specifically for only this design. Space-rated off-the-shelf

components can help to lower the cost function for the project.

Minimize the production cost: Once the design of the CUBIK is delivered to

TransOrbital, the company should be able to manufacture the kit at the minimum possible

cost. The design should be easily manufactured.

Use of Public domain codes: Software licensing introduces an unneeded expense

to the user. The design will use “freeware” such as Linux. In addition, the use of industry

standard I/O code will be used rather than designed.

Cost

Research and Development

Costs

Design Cost ($)

Production Expenses

Cost ($)

Amount of Software

Licenses which must be purchased

Cost ($)

MOE

Minimize

Maximize

Legend

Price ($)

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If ƒ denotes the cost function, there are three expressions. The three expressions

form are: ƒ Q A, ƒQB, ƒ Q (1/C) ∴ƒ Q(AB/C). The cost function for CUBIK is:

ƒ = k(AB/C) (1)

where k is the constant of proportionality, A is the research and development cost, B is

the production cost, and C is the cost of software licenses.

2.1.3 Simplicity Objectives The CUBIK system is designed for use by non-technical and amateur users.

Therefore, we want to make to make the system as simple as possible for users not

experienced with space technologies. There are three objectives for simplifying the use of

the CUBIK system for CubeSat, as shown in Figure 2.4.

Maximize Document Simplicity: Documents that the customer receives must be

clear and have a minimum of technical terms. Testing will be necessary to determine the

readability of the documents. Beta-testing will require non-technical volunteers to read

the documents. A first estimate of the simplicity before beta-testing will be done through

the use of the 80/20 document rule. Eighty percent of the document will actually contain

only 20% of the total vocabulary used.

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Figure 2.4 The sub-level objectives for maximizing simplicity.

Maximize User Interface Simplicity: The computer interface which the customer

uses to program the CUBIK system must be simple and easy to use. Beta-testing will be

the final test for simplicity and ease of use. Use of standard connection (serial port, USB,

etc.) and a basic Graphical User Interface (GUI) will help with ease of use.

Maximize Bus - to - Payload Interface Simplicity: Simplification of connections

between the CUBIK system and the user’s payload is important. Use of standard

connection types and use of generic I/O types are methods with help to achieve the goal

of simplifying the connection between the CUBIK bus and the payload. An estimation of

Simplicity

Clarity of Documents

Computer Programming Interface

Bus Computer and Payload Communications

80/20 Rule

Beta-test

I/O Function

User Computer Skill Level Required

GUI Interface

Manufacturing Difficulty

Time (s)

Process (#)

MOE

Minimize

Maximize

Legend

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the availability of the connection type will provide the first measure of effectiveness for

this goal. Beta-testing will provide the final measure of effectiveness.

2.2 Decision Hierarchy Design decisions are made by considering the importance of each objective in

relationship to all other objectives. This consideration is done by assigning a relative

weight of importance for each of the objectives in respect to all others. Then a matrix is

formed from those relative weights and normalized. The results of this process are shown

in Table 2.3.

Table 2.3 Decision Hierarchy Weights

Design Decision Weight Maximize power efficiency 0.0461 Minimize power consumption of bus---(Max power to payload) 0.0489 Minimize mass of bus 0.0752 Maximize additional payload volume 0.0106 Maximize stability 0.0889 Maximize communication range beyond LEO 0.0069 Minimize interference with other CubeSats/Rocket 0.0135 Maximize versatility for customer experiment 0.0841 Maximize computer efficiency 0.0320 Maximize structure integrity 0.0691 Minimize thermal effects 0.1553 Minimize radiation effects 0.1553 Minimize research and development costs 0.0412 Minimize productions costs 0.0676 Minimize amount of expensive software license purchases 0.0169 Maximize clarity of documents for user 0.0183 Maximize simplicity of computer interface to program the kit 0.0250 Maximize simplicity of interfaces between bus and payload 0.0451

The most mission critical of the performance goals are the minimization of

thermal and radiation effects. These objectives were given higher weights due to the

potential of radiation and thermal damage causing the destruction of the CUBIK kit. Cost

objective weights show minimization of production costs to be most critical. Simplicity

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goal weights show maximization of simplicity of interfaces between bus and payload to

be the most critical to the mission.

2.3 Summary The Value System Design process determines the objectives of the design and

establishes measures of effectiveness to evaluate the objectives. The objectives are to

maximize the performance of the design, to minimize its cost and to maximize its

simplicity. The development of each optimization is tabulated in an objective hierarchy

chart which is used to weigh the measures of effectiveness relative to each other. The

results are instrumental for developing the solutions presented in the next chapter, system

synthesis.

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Chapter 3 System Synthesis

3.1 Structure Because CUBIK is to be based on the CubeSat architecture, structural systems

synthesis aims to develop feasible alternatives for the mechanical elements not already

specified in the CubeSat design. Unspecified mechanical elements include: materials

used in construction, the internal structure of the bus, and any control or subsystem

mechanisms. Also, systems synthesis can be expanded to include mitigating problems

caused by the interactions between the structure and the environment. Interactions will

take the form of loading, vibrations, and thermal effects on the structure.

3.1.1 Material There are four constraints imposed on the choice of material. First, the material

should have thermal expansion properties similar to Aluminum 7075-T73, the material

used in the P-POD CubeSat launcher3. Similar expansion properties for the launcher and

satellite reduce the possibility of a launch failure due to reduced tolerances. Second, all

materials should be approved by NASA for use in space and allowed as acceptable

payload onboard any U.S., European, or Russian launch vehicle. Use of hazardous

materials will limit launch opportunities. Third, the material should be easily worked.

The CUBIK design objectives include maximizing the ease of manufacture for

TransOrbital, Inc. The fourth constraint pertains to minimizing the mass used for the

21

structure. Since the structural elements will likely account for the bulk of the allowable

bus mass of <1 kg, the material should have low density relative to strength.

Aluminum is the material recommended by Cal Poly, and used in every CubeSat

project which has been researched thus far. Aluminum is typically chosen to provide the

strength of a metal and to avoid the outgassing encountered by plastics in vacuum. Metals

such as aluminum are affected less by radiation exposure and temperature extremes than

are plastics and composites. The CUBIK team will use a metal, possibly Aluminum

7075-T73, for the structural components. There are other metals which have either low

density or a coefficient of thermal expansion comparable to that of Aluminum 7075-T73.

Three alternative choices of metal are stainless steels, magnesium alloys, and other

aluminum alloys10. Possible specific alloys include:

• Aluminum 7075 or 6061

• Magnesium AZ31

• Stainless Steel AISI 301

All three types of metal have properties which may meet the above constraints. All three

metals are also commonly used in space systems11.

3.1.2 Structural Elements

The CubeSat architecture designed at Stanford University uses four posts and six

panel faces as the outer structural elements. There are many variations of this

configuration, if chosen for the CUBIK bus. For example, the thickness of the panels is

not specified. Different sheet thicknesses are being considered for the panels to reduce

structure mass. Metal sheets are typically manufactured in a variety of thicknesses. In

22

particular, aluminum sheets are commercially available from some metal wholesalers in

any thickness from 0.01 to 0.25 inches12.

Other considerations include using a framework of panel strips, or perforated

panels in order to lower the over all mass while allowing more access for payload and bus

sensors. Similarly, there are alternatives to using solid posts. Square metal tubing and L-

shaped posts are being considered (See Figure 3.1).

Figure 3.1 Possible rail post configurations 3.1.3 Internal Structure Half the internal volume must be available to accommodate the payload. The

volume constraints dictate an efficient layout of electronics and other systems into the

available space. An alternative to designing a unique internal layout is adopting an

existing form factor. The PC/104 form factor has dimensions which make it an acceptable

option for use with CUBIK13.

Solid Square Tube L-shaped

23

Figure 3.2 PC/104 form factor dimensions13.

Internal components will require fasteners which secure them to the structure. The

structural components will also need to be fastened together. Options for fastening

include epoxies, bolts and nuts, screws, clips, and welds. Cost, ease of manufacture, mass

and suitability for use in space for each fastener type will be considered.

3.1.4 Control Elements, Bus Mechanisms, and Deployables The CubeSat specifications call for a deployment detection switch, separation

springs, and a “remove before flight” pin5. However, the particular locations of these

mechanical control elements are not specified. The locations may be adjusted to shift the

center of mass or to simplify the manufacturing process. Similarly, pin and switch

selection depend on mass, ease of manufacture, and cost. Alternatives exist for the

polymer material for the standoff foot required for the ends of the posts.

Polyoxymethylene (POM), with or without a Teflon coating, is being considered.

Other subsystems may employ mechanisms. The antennas may be deployable

rods, wire coils, or circuit boards affixed to the panels. Sensors may require mountings on

24

the surface of the panels or may need access holes through the panels in various

locations.

3.1.5 Thermal Effects Structural thermal regulation may be necessary. The choice of thermal regulation

depends on whether the satellite experiences any overheating or overcooling. In both

cases, thermal regulation by the structure will involve thermal conduction and thermal

radiation.

Thermal regulation schemes being considered for CUBIK include using the

structure as a heat sink and increasing the absorption of thermal radiation by interior

surfaces. Increasing the contact area of the bus/payload with the structure will provide for

more conduction of heat. Thermal radiation is the portion of the electromagnetic

spectrum extending from the wavelengths 0.1 to 100 µm. This band includes infrared and

visible wavelengths14. The interior surfaces may be painted black to increase absorption

in this band of the spectrum. Thermal analysis of the design is explored in later chapters.

3.2 Flight Computer

The flight computer on board is the hardware associated with the controls and the

functionality of CUBIK. During flight time, the computer is responsible for the power

control, communication and data storage. The flight computer detects the commands

from the ground station and passes it on to the flight computer interface. The flight

computer can be divided into two main components: the control unit and data storage.

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3.2.1 The Control Unit

Some primary responsibilities of the control unit are to detect commands from the

ground station, control power regulation, and to format data for downlink and output

them to the communication subsystem. The CU can be made of either a microprocessor

or a micro controller. The possible choices for these are:

• Motorola

• Hitachi

• Intel

Our choice of either a microprocessor or a micro controller will be bounded by some

constraints. Some of which include:

• It should operate from -40 to 150 degree Celsius.

• It should have a PROM (Programmable Read Only Memory)

• It should have minimum wear and tear if exposed to space radiation.

The CU should be able to initiate the system after deployment. Hence, the PROM is

programmed to commence the start sequence.

3.2.2 The Data Storage

The data storage is yet another important aspect of the flight computer. The type

of processor or controller used will affect the type of memory used for the BUS.

Microcontrollers have a ROM (Read Only Memory) which can be programmed to

perform a specific function. We then use an external memory chip to store the data from

the experimental payload, should we require storing it on board and not down-linking

back to the ground station. The decision on the selection of the memory will be bounded

by certain constraints. The type of memory chosen should:

26

• Be operational in the temperature range of -55 to 120 degrees Celsius

• Consume low power

• Be able to handle sufficiently large data and

• Have minimum wear and tear when exposed to space radiations.

The above restrictions limit the brands of memory which can be employed on CUBIK.

3.3 Flight Computer Software

The flight computer software (FCS) loaded on the on board flight computer

coordinates and commands every action of the satellite. The main component of the FCS

is the operating system (OS). The OS is the root of the system and gives command to all

the subsystems of the spacecraft, namely the communications, ADCS, power, and

payload. All these subsystems are controlled by the FCS through drivers. The OS is also

supposed to handle real time operations, which requires the OS to be a Real Time

Operating System (RTOS).

3.3.1 The Real Time Operating System

The RTOS is the interface between all the subsystems of the spacecraft and the

Central Processing Unit (CPU). The possibilities for the RTOS are numerous and depend

on the architecture of the CPU. The alternatives for the CPU are either to use a micro

controller or a microprocessor. The most widely known Open Source RTOS is Linux and

is available for both the microprocessor and the micro controller. Other RTOS exist, such

as QNX for 32 bits microprocessors. These RTOS are Open Source, which means that

their source code is available to the customer and can be changed according to the

27

customer’s need. This type of license is called GNU Public License (GPL) under which

the software is distributed.

The term used to describe the use of RTOS for non-desktop or server use is

embedded technology. All the previously mentioned RTOS have embedded technology

software derived from their original product.

3.3.2 The Micro Controller Case

The micro controller, as presented in section 3.1.1, is a CPU without memory

management unit (MMU), thus the OS has to be designed specially for this case. Many

alternative RTOS exist for specific micro controllers so that their choice depends on the

actual type and brand of the micro controller. However, Linux has an extended embedded

technology and even a Linux RTOS written for micro controllers, uClinux. Both of these

Linux systems allow a wide range of micro controllers. The specially written Linux OS is

not originally real time, however a software “add-on” can be made to the OS to make it a

RTOS.

3.3.3 The Microprocessor Case

The microprocessor is a CPU more capable than the micro controller, it exists a

large variety of microprocessors with different architecture that the OS have to handle

differently. However, a number of OS are real RTOS. These include Linux for every

type of architecture possible, ranging from Intel to Alpha, RISC, and SPARC. The other

possibility is QNX.

3.3.5 The Drivers for Subsystem Devices

The FCS will manage all of the subsystems, so each piece of equipment in a

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subsystem must be controllable by the FCS. The FCS commands the different devices in

the subsystems through drivers, which are little programs written and installed in the FCS

that convert the commands from the FCS to a language understood by the devices. These

drivers are specific to a particular device. To write these drivers' codes, different

languages can be used, such as C/C++ or ADA. The only requirement, for simplicity and

efficiency, is that the languages have to be Object Oriented, since only such languages

have tools performance enough to permit the creation of such complicated software.

Most of the Open Source RTOS have an integrated C/C++ compiler, so there is no need

to add another compiler if this language is chosen.

3.4 Communication

An important component in satellite design is the communications system. The

main aim of a communication system is its ability to receive information from a ground

station and transmit data or information to a ground station. The Cubic communication

system will consist of transponder and an antenna.

The communication system will communicate using armature radio frequencies.

Armature radio frequencies are used because they are free and easy accessibility. There

exists commonly available equipment supporting these frequencies with a huge network

supported by well-maintained and equipped ground stations.

3.4.1 Transponders

Transponder is a device that combines both a receiver and a transmitter on one

board. Due to their varied functional requirements, most transponders must be custom-

29

designed to be able to function according to what the user wants. Two major types

considered are linear and FM transponders. The major companies producing these

transponder chips include Motorola and Alinco.

Satellites using linear transponders receive a specific range of frequencies

(typically 40 - 100 kHz) in one band, convert them to another band using a mixing

process and amplify the converted signal for transmission back to Earth. The signal

received by the satellite referred to as the uplink signal and the signal transmitted to earth

is referred to as the downlink signal. Advantages of linear transponders include:

• Linear mode allows more than one station to use the transponder at a time.

Meaning more probability to get a stronger signal therefore better reception

• Little power is used when the transponder is not in use. We cannot always turn

transponder on and off continuously, so this property is important

• Doppler shift is handled by the ground station which means we have to worry less

about correction to signal transmission.

Disadvantages of linear transponders include:

• More complex ground stations than with FM and are more costly in some cases

• Congestion might occur on transponder since it allows more than one station to

used the transponder at one time

Satellites using FM transponders use the concept of frequency modulation to

transmit and receive information back and forth from the earth to the satellite.

Advantages of FM transponders include:

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• New users can be recruited using their existing equipment.

• A transponder easier to construct with readily available chip sets.

• Transponder power is concentrated on one user meaning better signal quality

Disadvantages of FM transponders include:

• Only one station can use the satellite at any time.

• The strongest uplink signal will capture the satellite.

• Full transmitter power is used at all times (unless turned off).

• Higher average power consumption than a linear transponder.

The choice of a transponder will depend on what experiments we will be

supporting on board the satellite to the user. It will also very according to the type of

antenna we use and the amount of overall power we can allocate to this device.

3.4.2 GPS receivers

The function of the GPS receiver is to determine the position of the satellite in

space and transmit this to user to be able to determine precise position. There are

numerous GPS receivers that could be used but the ones with the lower power

consumptions.

GPS receivers with low power consumptions include Royal Tech REB-2100,

REB-2000 and RGM-3000. The advantages of REB-2100 and (REB-2000) are twelve

parallel channels for data transfer, 0.1 seconds reacquisition time and compact size (30 X

30 X 8 mm). The disadvantage is that these require an external antenna to be mounted

externally on the satellite. The advantages of REB-3000 are twelve parallel channels,

built in path antenna and high sensitivity to GPS satellite signal and 20% lower power

31

consumption than the other two. The disadvantage compared to the other two is a $10

higher price than REB-2100 and REB-2000 ($500).

The main computer could do the job of GPS receiver system providing less

accurate information. A GPS receiver system is important with a disadvantage being its

cost. In conclusion a decision has to be done on how accurate we need our satellite

position with respect to how much budget is available. In other words how much we want

to maximize our performance versus minimizing our cost.

3.4.4 Antennas

An antenna is any structure or device used to collect or radiate electromagnetic

waves. The antenna plays an important role in working the satellite. It is the first part of

your communication system that will receive the signal and the last one to transmit it.

The types of antennas that could be used are circular polarized antennas and single plane

antenna. Some of the brands making good antennas for satellite use are KLM, M2,

Hygain and Cushcraft.

Circular antennas suitable for CUBIK come in two dimensions: 2m or 70cm. The

antennas are phased so that the signal normally rotates to the right, meaning the antenna

is right hand polarized. The circular polarized antennas are about $500, however the

antenna transmit and receive signal with very high quality.

Single plane antennas can mounted vertically or horizontally when operating the

satellite. There are two main types, quarter wavelength (l/4) and half wavelength (l/2).

The wavelength refers to the size of the antenna. A half wavelength antenna has better

signal reception than a quarter wavelength antenna. However consumes more power than

32

a quarter wavelength antenna. These antennas are very cheap and are easier to deploy

once in space than a circular antenna. They also take much less space and are smaller in

size. The problem with those antennas is if the satellite is spinning the signals will fade in

and out.

In terms of the number of antennas needed on the satellite we can have one

antenna for uplink of information and two for downlink of information. Another way is to

have two antennas for uplink of information and two for downlink of information. The

reason for having two antennas for downlink is because transmitting consumes more

power. One antenna cannot do the job efficiently and it would require more time to send

information with one antenna rather than two. Having two antennas for uplink would

make the receiving more efficient with the drawback being the extra weight and cost.

There are two deployment mechanisms possible for the antennas available on the

CUBIC are a crucial part in the successful of our satellite. If the deployment does not

work the satellite will have no communication capability. One solution is to wrap up the

antenna around the satellite structure and release the antenna by a spring mechanism after

the satellite has been launched into space and receives the ground command to release the

antenna. This mechanism is used for circular antennas. Another deployment mechanism

is a nylon line attached to the tip of the communication antenna, then a nichrome wire is

winded in a spiral and the nylon line is threaded into it. The nichrome wire is then heated

by an electric, which this causes the nylon wire to melt and release the antenna. The

communication antenna is then deployed by its own elasticity. This mechanism is used

for single plane antennas. The advantage of this is the possibility of having the wire stuck

is less than the first method since we have the antenna inside the cube. The disadvantage

is that a smaller antenna has to be used since it must be stored inside the cube.

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3.5 Attitude Determination and Control System

The attitude determination and control system determines the satellite’s

orientation with respect to some external object and keeps the satellite in a desired

orientation. Satellites can have attitude determination without control. In a case of a

satellite without attitude control, the determination acquires the satellite’s orientation and

can control the payload and communications systems to operate at times where the

orientation is correct for useful data collection and transmission. The two systems are

described here separately after a discussion of satellite attitude stability.

3.5.1 Attitude Stability A satellite can have three forms of attitude stability in orbit. The first form is no

stability. When a satellite has no attitude stability it is free to rotate in orbit on all three

axes at once. Free rotation is called tumbling. When tumbling, there is a small possibility

that the tumble will prevent alignment of sensors and communication to a useful

orientation. Some missions and communication systems may not require a particular

orientation and therefore will not need ADCS. For example a mission which is to

transport ashes into space will not need ADCS in the satellite carrying the ashes.

The second form of attitude stability is single-axis stability. Single-axis stability

requires that one axis be maintained at a particular orientation. The satellite can spin

around that axis, which is a method to maintain the stability. A mission requiring single-

axis stability may be one with a sensor located on the face perpendicular to the spin

pointed at a particular object. A mission requiring changes of orientation can destabilize

34

the one axis of stability in a controlled manner to allow it to change orientation to the

desired direction. Some methods of one axis stability, such as gravity gradient control,

will not allow this. With gravity gradient stability, the single-axis of stability is forced to

be perpendicular to the earth’s surface, preventing reorientation of that axis to another

position, for example perpendicular to earth’s surface.

The third form of attitude stability is three axes stability. Three axes stability

requires that all three orthogonal axes be maintained in orientation with respect to an

external object (say the earth surface). Three axes stability can be required by a mission

that uses a sensor to take continuous real time readings. Some methods of three axes

stability will allow the satellite to be reoriented to any desired orientation.

Figure 3.3: The three forms of attitude stability.

3.5.2 Causes of Attitude Instability

There are several forces which work to destabilize the attitude of satellites15.

Electromagnetic flux as the satellite travels through the magnetic flow lines can cause

35

torques if the satellite contains circuitry or a relatively large ferro-magnetic component.

In space industry situations, electromagnetic flux would be resolved by design. In the

case of selling the CubeSat CUBIK kit for amateurs to put in their payload,

electromagnetic flux becomes a consideration for ADCS.

Rotating machinery (pumps, tape recorders, rotating sensors, etc.) create torques

that can perturb both stability and pointing accuracy. Liquid sloshing creates torques due

to fluid motion and variation of center-of-mass location. Also uncertain will be the final

center-of-gravity due to the placement of the payload installed by the customer who

purchases the CubeSat CUBIK kit. Volumetrically off-centered center of gravity can

create unwanted torques during attitude control.

Dynamics of flexible bodies can also cause oscillatory responses at bending

frequencies. Third party payload experiments containing flexible and bending

components will cause vibrations that can limit the control response frequencies of the

attitude control system. Since the frequencies of the payload are unknown during design,

the documentation must include limits on the payload natural frequencies.

Thermal shocks on flexible appendages cause attitude disturbances when

entering/leaving eclipse. For the CUBIK design, thermal shocks are only a consideration

if a gravity gradient boom is used or if a third party experiment has an extending

appendage.

The various causes of attitude instability on the satellite can only be fully

determined with full knowledge of the third party payload. Assumptions must be made

and a constraint must be set on the payload, depending on the final type of ADCS used.

The assumptions and limitations are determined in the system analysis phase.

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3.5.3 Attitude Determination

Attitude determination is the method of ascertaining the satellite’s orientation in

space with respect to one or more outside objects. Sensors are used to determine the

attitude. RFP requirements state that at least one Sun sensor be included in the CUBIK

design. The Sun sensor alone is insufficient to determine the attitude of the satellite.

Considering a tumbling satellite, the Sun sensor will only tell if the sensor is pointing to

the sun at any moment, the orientation of all orthogonal faces remains unknown. The

position of any other payload sensors or communication transceivers on the orthogonal

faces remains unknown. For some types of missions tumbling is not a concern. For

example a mission to study the Sun would need its sensors on the Sun sensor face.

Missions studying the Earth, Moon or space will require more attitude determination.

Other sensors (See Table 3.1) would need to be used.

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Table 3.1: Attitude Determination Sensors15

Sensor Sensor Determination Sensor Limitation

Sun Sensor Angle of Sun to Sensor Requires unobstructed view, Loss of data during eclipse period in LEO

Star Sensor Mode 1: Scanner: Determines attitude from multiple crossing of star in field of view. Used in spinning satellite.

Mode 2: Tracker: Tracks star to determine attitude. Used in 3 axis stabilized satellite

Sensor is blinded by Sun, Moon, Earth and Planets that may cross the field of view.

Horizon Sensor

Infrared signature of Earth horizon:

Type 1: Pipper: Narrow field of view uses angles for determination. Used on spinning satellite.

Type 2: Scanning: Rotates lens or mirror to get angles for determination. Used on spinning and non-spinning satellites.

Requires clear field of view. Need to account for seasonal variations in LEO

Magnetometer Measures size and direction of Earth magnetic field for determination

Lower accuracy due to changes in magnetic field. Operation must be cycled when using magnetic torques for control.

GPS Receiver Uses difference of position of two (or more) receivers to determine attitude.

Accuracy related to distance between receivers.

Gyroscope Measures speed and angle of rotation for determination

Requires other sensors for attitude determination.

Combining sensor types can mitigate or eliminate limitations. Combining a Sun

sensor with a Horizon sensor and a Star sensor will give high accuracy (< 0.1o)10 attitude

determination and maintain determination during eclipse. Combining sensors increases

cost. Optimization analysis will need to determine the optimal configuration for an

assumed payload requirement.

3.5.4 Attitude Control

The pointing options for the missions determine the method of attitude control.

Third party payload requirements are unknown and are assumed to be restrictive. Cost

factors of the attitude control system used are constrained by the given cost limit.

38

Optimization during system analysis considers the options available vs. the cost. Options

available have differing attitude control characteristics. For example, gravity gradient

stabilization limits the two faces perpendicular to the boom to pointing at or directly

away from the Earth. Propellant for attitude control of the satellite is not allowed in the

Request For Proposal. Those options requiring propellant are not considered. The

available options and characteristics are in Table 3.2.

Table 3.2: Attitude Control Options for CubeSat CUBIK System15

Control Method

Pointing Options

Attitude Maneuverability

Typical Accuracy

Lifetime Limits

None None None +/-180 deg(3 axis) None Gravity-gradient Earth local

vertical only Very limited +/- 5 deg (2 axis) None

Passive magnetic North/South only Very limited +/- 5 deg (2 axis) None Pure Spin Stabilization

Inertially fixed in any direction. Repoint with precession maneuvers.

High power consumption by magnetic torques to move stiff momentum vector

+/-0.1 to +/- 1 deg in two axis (proportional to spin rate)

None

Bias Momentum (1 wheel)

Local vertical pointing

Momentum vector of the bias wheel prefers to stay normal to orbit plane, constraining yaw maneuver

+/-0.1 to +/- 1 deg Life of sensor and Wheel Bearings

Zero Momentum (3 wheels)

No constraint No constraint +/-0.001 to +/- 1 deg

Life of sensor and Wheel Bearings

Zero Momentum CMG

No constraint No constraint High rates possible

+/-0.001 to +/- 1 deg

Life of sensor and Wheel Bearings

Control actuators are determined by the method used and can be combined to

achieve greater control and/or accuracy. One option is to use no control. No control will

allow the satellite to spin at any rate on all axes simultaneously. Third party payload type

will be limited to those types not requiring any attitude stability. For example omni

directional transmission or local study of electromagnetic field strength does not require

attitude stability.

39

A gravity gradient boom will passively point one axis of the satellite toward the

Earth center of gravity. The boom requires no power beyond the deployment of the

boom. The boom can stabilize both pointing to and pointing away from the Earth. The

boom will interact with the satellite structure after deployment.

Reaction wheels are torque motors with high inertia rotors. Momentum wheels are

reaction wheels with a nominal (minimum) spin rate to provide a nearly constant angular

momentum. Angular momentum provides gyroscopic stiffness to two axes while the

motor torque controls the pointing around the third axis. Wheels have a saturation speed.

The restriction of a saturation speed means cyclic disturbances can cause the wheel to

reach a maximum speed, after which there is no more control over the cyclic disturbance.

Usually thrusters are used to de-spin the wheels. Thrusters are not an option allowed for

CUBIK in the RFP. Magnetic torques may be used to achieve the same result if analysis

shows that spin saturation may be a problem in the required 1 year lifespan. Wheels can

provide one or three axes of stability depending on how many wheels (one or three non-

coplanar) are used. Reaction wheels may be used instead of propellant to create a spin-

stabilized satellite. Reaction and momentum wheels interact with the spacecraft

structure, the power system and the flight computer.

Control-moment gyros are momentum wheels with a high rate of spin. In

comparison to reaction and momentum wheels, control-moment gyros are high cost and

high mass. Due to cost and mass constraints they may not be used for three-axes stability.

One gyro could be used to create a spin-stabilized satellite. Control-moment gyros

interact with the spacecraft structure, the power system and the flight computer.

40

Magnetic torques use magnetic coils or electromagnets to produce dipole

moments that react with the Earth’s magnetic fields. Magnetic torques are effective in

LEO where the magnetic field strength is strongest. Magnetic torques can be used for

attitude control or for de-spinning of momentum/reaction wheels. Magnetic torques

interact with the spacecraft structure, the power system, the flight computer, and all

electronics including payload.

3.5.5 ADCS Best Options Numerous sensors and actuators result in a large number of possible combinations

for attitude determination and control. The TransOrbital RFP requirement of no back-ups

or redundancy in systems and structural deployment only with permission of the

requestor in conjunction with the lack of known requirements of attitude determination

and control for the third party payload results in the assumptions of versatility in the

CUBIK system and manageability of the number of options. The assumptions are: no

boom deployment to advert necessity of permission for such; three wheels or three

magnetic torques maximum, leaving out redundant and backup systems; no more than

three actuators of any combination; no star sensors due to ability to be blinded by other

emissive/reflective objects.

The assumptions result in the combinations of: Sun and Horizon; Sun and

Magnetometer; Sun and Horizon and Magnetometer sensors for attitude determination.

The possible options for attitude control are in three groups: no stability, one-axis

stability, and three-axes stability. The option of no stability does not require actuators.

One-axis stability can use one gyro to create spin stabilization, one momentum/reaction

wheel to stabilize the axis or two to create spin stabilization, one magnetic torque to align

41

to North/South attitude axis or two torques to create spin stabilization. Three-axes

stabilization can use three non-planar momentum/reaction wheels, three orthogonal

magnetic torques, or a combination of three momentum/reaction wheels and magnetic

torques.

3.5.6 ADCS Summary

The options which will be used will be determined by system analysis in terms of

performance, cost and mass. Performance will be secondary to cost and mass. Given the

constraints on mass and cost, one alternative of CUBIK system may feature only attitude

determination without control.

3.6 Power

Power for the CubeSat will be generated by photovoltaic cells, also known as

solar cells. The request for proposal specifically asked for the use of solar panels

covering at least five of the six sides of the satellite. Solar power is a good choice due to

its reliability and relatively low cost. The choice of exactly what brand and type of solar

cell to be used will be made based on a balance between efficiency and cost.

3.6.1 Power Generation

Power for the CubeSat will be generated by photovoltaic cells, also known as

solar cells. The request for proposal specifically asked for the use of solar panels

covering at least five of the six sides of the satellite. Solar power is a good choice due to

its reliability and relatively low cost. The choice of exactly what brand and type of solar

cell to be used will be made based on a balance between efficiency and cost.

42

3.6.2 Power Storage

The solar cells will only be producing power for a portion of the orbit period due

to the satellite moving in and out of direct sunlight. For this reason the satellite will need

to be able to store energy for use during the periods when it is in eclipse. Due to size

constraints, the satellite is limited in the types of energy storage it can accommodate.

Small flywheels or batteries are the best alternatives due to their size. A flywheel stores

kinetic energy by being spun up by a motor, and later releases that energy by spinning

down and generating power through the motor. A battery uses chemical reactions to

release energy and recharges by reversing the process. Batteries are reliable and have

been used extensively in spacecraft. Three types of batteries being considered are, nickel

cadmium (NiCd), nickel metal hydrides, and lithium ion.

Flywheels have been used to reliably store energy, and are becoming more

popular for use in small satellites. A possible drawback is that flywheels create a torque,

but this torque could possibly be incorporated into the attitude determination and control

system (ADCS). Size is another drawback to the use of flywheels, but small flywheels

designed specifically for satellite use do exist.

Nickel cadmium (NiCd) batteries are popular, and depending on the battery can

endure up to 900 charge/discharge cycles. Nickel cadmium batteries are endothermic

while charging, and could be used as part of the thermal control when the satellite is in

direct sunlight. Nickel cadmium batteries have a lower power density than nickel metal

hydrides and lithium ion batteries. Power density is a measure of amount of energy

stored per kilogram. One major drawback to the use of NiCd batteries is a memory

effect, meaning that if they are recharged before fully discharging, they will “remember”

43

the level to which they last discharged and will not discharge past that point later.

Another drawback is that cadmium is a toxic material and could pose a health risk if the

battery were to rupture.

Nickel metal hydride batteries have approximately %50 more storage capacity

than NiCd batteries due to a higher power density. Nickel metal hydride batteries are

about equal to NiCd batteries in the amount of charge/discharge cycles they can handle.

Unlike NiCd batteries, nickel metal hydride batteries have a minimal memory affect.

They are also exothermic during charging which could pose a problem by creating extra

heat during the period when the temperatures of the satellite are at their highest. Nickel

metal hydrides also tend to have a high self discharge rate, sometimes as high as %5 of

the total capacity lost per day.

The third type of batteries being considered are lithium ion batteries. These

batteries have up to %50 more storage capacity than even nickel metal hydrides. The

extra storage capacity is due to the extremely high power density of lithium. Lithium ion

batteries tend to have a longer battery life than NiCd or nickel metal hydride batteries.

They also can withstand up to 1200 charge/discharge cycles, and have a minimal memory

effect. The only drawback is that lithium ion batteries have a high internal resistance,

which means that they cannot deliver high currents. Power requirements will have to be

determined before it is known if a high resistance will pose a problem.

3.6.3 Power Regulation and Distribution

The power harnessed by the solar cells must be regulated before it can be

distributed among the other subsystems and the batteries. When power is regulated the

voltage is stepped up or down to the voltage required by the load. Proper power

44

regulation can also prevent problems such as fluctuating voltage, voltage sag, and surge

problems. Micrel, National Semiconductor, and Ricoh all have lists of potentially usable

voltage regulators in their line of integrated circuits. The cost of these integrated circuits

is minimal, typically less than fifty cents, allowing selection based solely on the needs of

the satellite. Selection of a power regulation integrated circuit will be postponed until the

details of the power supply and power requirements have been defined.

3.7 Documentation

In order to make CUBIK a marketable product, its specifications must be

documented. Clear, concise, and easy to understand directions for manufacturing and

using the kit must be provided. The documentation falls into two categories: instructions

for TransOrbital, Inc. and instructions for customers of the kit. TransOrbital needs

documentation to the extent that the company can readily produce the subsystems for

sale. Users of the kit must be provided with an easily understood operating manual.

Alternatives for documenting the specifications and instructions for both TransOrbital,

Inc. and the CUBIK users need to be generated. Some examples of ways to document

include:

• Microsoft Word document, PDF file, or other format

• Movie on video, compact disc, or DVD

• Paper copy of operating manual

• Hard copy on compact disc of operating manual

The above examples can be used alone or in combination with others.

45

3.7.1 Documentation for TransOrbital, Inc.

The documentation for TransOrbital, Inc. must be converted to a form which can

be easily understood and utilized. TransOrbital could be provided with a Microsoft Word

document on CD or a hard copy on paper. These three forms of documentation would

have to include assembly drawings, results of all the testing such as thermal and

vibration, and the source code for all software elements. Using a CD could be a feasible

solution because assembly drawings would be easy to download from a compact disc. A

Microsoft Word document including all drawings, codes, and test results is also feasible

because Word is a standard program. The important factor is that the documentation for

TransOrbital must be detailed enough so the company can manufacture the subsystems

for sale. Whichever type of documentation accomplishes this goal most efficiently will

be a viable option.

3.7.2 Documentation for the User Customers of the kit must be provided with an operating manual. The manual

should come in a universal form or forms, which most users would be able to exploit.

The user could be provided with a compact disc which would include the dimensioned

physical drawings, lists of materials used, a frequency band analysis, electronic

subsystem schematic drawings and list of components, and software interface and

operating specifications. All this data could be formatted using Microsoft Word and

Excel. Microsoft programs are feasible because they are universal, allowing for a

majority of users to view the operating manual with no difficulty.

46

A video could be provided for the users to show them how to work the interfaces

or install the payload. Instead of producing the operating manual on a compact disc, it

could be done in the conventional way as a paper copy. The types of documentation

mentioned above are all feasible solutions. The critical aspects of the documentation to

the user include easy readability, non-complicated wording, and minimal use of

technological details. The form or combination of forms of documentation which

achieve these objectives efficiently will be a feasible solution.

3.8 Summary

The System Synthesis chapter reviewed the six main subsystems of CUBIK, in

addition to documentation. The chapter described alternative methods for achieving the

objectives of each subsystem. In relation to the objectives and constraints, the viability of

the alternatives is discussed. In conclusion, System Synthesis presents a bunch of

feasible solutions for each subsystem and attends to the feasibility issues of each

alternative.

47

Chapter 4: System Analysis

System analysis accomplishes the task of evaluating different alternatives with

respect to three design models: Minimum Cost, Maximum Performance, and Maximum

Simplicity. Each concept is a maximization of one of CUBIK project objectives.

In the following sections, the analysis is divided by subsystems. The alternative

solutions for each subsystem are examined and analyzed with respect to each design

model. A “best choice” for each design concept will be chosen. Picking a “best choice”

provides a simple way to model the system analysis. The three design concepts are not

final designs; rather they are basic concepts which are subsequently used to provide

structure to the modeling and analysis presented in this chapter.

4.1 Structure The structure subsystem goals are to provide structural support for the bus

components, control elements, and payload and mitigating hazardous effects from the

environment such as shocks and thermal radiation. The main structure subsystem

components are: the structural elements, internal structures, and control elements. The

system analysis for the structure subsystem compares the alternatives in relation to our

three design models: minimum cost, maximum performance, and maximum simplicity.

Alternatives presented in Chapter 3 which violate design constraints are abandoned at this

stage.

48

4.1.1 Material The three materials considered for the external structural elements were

aluminum, magnesium and stainless steel alloys. The material properties shown in Table

4.1 were compared with Aluminum-7075, the structure material for the P-POD launcher.

Magnesium alloys and stainless steel alloys were found to have unacceptable material

properties which make aluminum the material of choice. In particular, the thermal

expansion coefficient, a, for stainless steels was low relative to Al-7075. This introduces

the possibility of reduced tolerances between the P-POD launcher and the CUBIK rails,

leading to ejection failures. Stainless steels also unsuitable because they are relatively

massive compared with the aluminum and magnesium alloys. Research uncovered that

magnesium is relatively difficult to machine. Fire suppression precautions must be taken

during machining magnesium because of the possibility of ignition19. This precaution

violates the objective to produce a design which is simple to manufacture.

Table 4.1 Material Properties of the Alloys Considered for the Structural Elements

Material Density (kg/m3)

Coefficient of Thermal

Expansion (µm/m-°C)

Relative Availability

Relative Machinability

Aluminum 2700 23.6 Excellent Excellent

Magnesium 1770 26.0 Poor Poor

Stainless Steel 7900 16.6 Excellent Good

All three design models will use aluminum alloy external structural elements.

Aluminum alloys are available in many forms, from numerous suppliers. Aluminum

alloys also have excellent machinability.

49

4.1.2 Structural Elements As shown in Chapter 3, there are a variety of structural component shapes.

Preliminary design models are based on the CubeSat schematic provided by Stanford

University and prototypes created by other CubeSat project teams. Those designs consist

of posts and panels.

In general, less fabrication of components results in lower production costs. For

the low cost model, the structural elements would consist of solid sheet aluminum panels

attached to extruded, solid square aluminum posts.

The maximum performance model employs square tube posts in order to lower

mass and increase strength. Similarly, the panels would not be solid sheets. Panels would

be perforated and would provide the minimum support necessary.

Maximum simplicity is achieved by using structure elements that are easy to

fabricate and assemble, and minimally interfere with other subsystems. The maximum

simplicity model employs L-shaped posts which allow fitting unmodified PC104 boards

inside the structure. This post design reduces structural mass compared to using solid

posts. Panels would be perforated sheets.

4.1.3 Internal Structure The internal structure alternatives are to use the PC104 form factor or to use a

custom design. The advantages of using the PC104 form factor include the ability to

purchase rather than manufacture electronics boards. Publicly available PC104

specifications reduce the new for creating documentation.

The minimum cost and maximum simplicity models use the PC104 form factor.

The maximum performance model uses a custom designed internal structure. Custom

50

designed internal structures require more engineering but will allow the internal volume

to be used more efficiently.

4.2 Flight Computer

The flight computer subsystem is the brain of the CUBIK bus. The flight

computer optimizes and regulates the performance of the electronics in the bus.

Communications between subsystems and operational synchronization between

electronic components is maintained by the computer, the common interface among all of

the bus subsystems. The mission is controlled by the flight computer. In particular, the

computer interprets commands from the ground station. The computer also stores all

experimental data and performs all necessary computations.

The flight computer is comprised of three main units: a storage unit, a processing

unit and a control unit. Figure 4.1, shows the typical interactions of the three units.

Figure 4.1 A generalized representation of the flight computer.

The other subsystems send signals to the flight computer and the computer

interprets the signal and sends an appropriate response either back to the subsystem or to

the appropriate unit.

Processing Unit

Control Unit

Storage Unit

51

4.2.1 The Storage Unit

The experimental data along with the critical data needed for the functioning of

CUBIK are stored in the storage unit. The storage unit is only accessible through the

flight computer and the flight computer software. Storage needs depend on the

experimental payload and the volume of data that the payload collects. A greater amount

of memory means a larger consumption of power.

Depending on the storage needs, the following types of memory devices can be

used: RAM/ROM, flash memory, or memory sticks. Table 4.2 summarizes memory

characteristics.

Table 4.2 Characteristics of Memory Devices Memory Device Advantages Disadvantages

RAM or ROM Easy to install and operate Inexpensive

Unreliable at extreme temperatures

Flash Memory Easy to install and operate Unreliable at extreme temperatures

Memory Sticks Reliable at extreme temperatures Expensive Require additional computer BUS

4.2.2 The Processing Unit

The processing unit is the most important component of the flight computer. All

the processing for the experimental data and the processing for the commands is done by

the processing unit.

A choice of the processor for the processing unit will depend on the amount of

computation that the payload experiment needs to perform. For a fairly simple

experimental payload a processor with relatively low processing speed will work just

fine, but on the other hand, for a payload which needs pretty complex and fairly large

amount of computation, a processor with faster computing capabilities is needed.

52

Again, the power supply limits the kind of processor that can be used for the

CubeSat. A faster processor means large power consumption. The following table helps

you to understand the differences between the different processors that Intel has

introduced over the years.

Table 4.3 Intel Processors Operating Characteristics CPU Date Transistors Microns Clock

speed Data width

MIPS

8080 1974 6,000 6 2 MHz 8 bits 0.64 8088 1979 29,000 3 5 MHz 16 bits

8-bit bus

0.33

80286 1982 134,000 1.5 6 MHz 16 bits 1 80386 1985 275,000 1.5 16

MHz 32 bits 5

80486 1989 1,200,000 1 25 MHz

32 bits 20

Pentium 1993 3,100,000 0.8 60 MHz

32 bits 64-bit bus

100

Pentium II

1997 7,500,000 0.35 233 MHz

32 bits 64-bit bus

~300

Pentium III

1999 9,500,000 0.25 450 MHz

32 bits 64-bit bus

~510

Pentium 4

2000 42,000,000 0.18 1.5 GHz

32 bits 64-

bitbus

~1,700

4.2.3 The Control Unit

The control unit communicates with the electronics subsystems, sending and

receiving computer commands from the ground station and subsystems. The control unit

also manages the power among the subsystems by distributing the power among them

and selecting the voltage source to draw power from.

53

4.3 Flight Computer Software

The Flight Computer Software (FCS) is a critical subsystem of the spacecraft

since it coordinates and commands every action of the spacecraft. Choice of FCS affects

the performance and reliability of the spacecraft. The FCS is composed of an OS and

software that handles communications, telemetry and spacecraft device drivers.

4.3.1 The Operating System The OS should be a Real Time Operating System (RTOS). System synthesis

alternatives for the microprocessor architecture are: Linux, an open source UNIX based

system, and QNX. Microcontroller architectures use uClinux. The characteristics of the

operating systems are shown in Table 4.4. Windows CE, which is a Windows RTOS

designed for embedded technologies is included for comparison purposes17. Due to

software licensing considerations, Windows CE is not a suitable OS for the low cost

model of CUBIK.

Table 4.4 Operating System Alternatives and Specifications

Operating System

Supported Architecture

License Real Time Integrated Compiler

Linux x86, Alpha SPARC, PowerPC

GPL Yes C/C++

QNX x86, MIPS PowerPC

GPL Sources Partly available

Yes None

uClinux Microcontrollers Hitachi, Motorola

GPL Yes with patch C/C++

Windows CE ARM, MIPS, SHx, x86

Commercial Yes Platform Builder

Linux can be easily modified to meet the performance requirements of CUBIK.

Specifically, the OS core can be lightened to contain only the necessary software

components, such as the OS kernel, the integrated C/C++ compiler and necessary

54

functions. Linux supports a wide range of processor architectures, which makes it a

suitable choice for different designs and performance levels. Furthermore, Linux is

distributed under a GPL license, not a commercial license, which makes it an inexpensive

OS. As a result of its licensing and flexibility, Linux fits the requirements of both the low

cost design model and high performance design model.

The QNX OS has the same characteristics and performance as the Linux OS with

the exception that only part of the source code is open source. Moreover, unlike Linux,

QNX does not have an integrated compiler. An integrated compiler is helpful for

software development and helps ensure software compatibility with the hardware. The

compiler for QNX is commercially available. Due to the additional cost of the QNX

compiler, QNX is only suitable for the high performance model.

uClinux is an OS similar to Linux but specially designed to work on

microcontroller architectures. This OS can be modified to operate as a RTOS by using

software available on the internet. This OS is the choice for all the models which use

microcontrollers.

4.3.2 The Programming Language The FCS uses an assembler to change the programming language into assembly

code, which is then read and understood by the processor. The programming language is

used to write software to handle complicated tasks such as communication decoding,

telemetry formatting and spacecraft devices handling. Object oriented languages have

tools to achieve these tasks. The programming language alternatives considered are

C/C++ and ADA. The languages have similar characteristics as well as performance;

however C/C++ is more popular among developers and amateur programmers. For this

reason, C/C++ is suitable for all the three design concepts.

55

4.4 Communications The communication subsystem goals are: receive information from the ground

station, and transmit data from the satellite to ground. The main communication

subsystem components are transponders, antennas and GPS receivers. A general

representation of the communication system and interactions is shown in Figure 4.2.

Figure 4.2 The communication subsystem components and interactions.

4.4.1 Transponders This section presents the advantages and disadvantages for each type of

transponder alternative relative to the three design concepts. The two choices under

consideration are FM transponders and linear transponders. Table 4.5 summarizes the

two transponders.

Table 4.5 Transponder Characteristic Transponder Number of

Stations Receiving Signal

Power (mW)

Production Cost ($)

FM 1 at a time 2 Chips readily available

100

Linear 20+ 3 May require special order/design

20+

Flight Computer

Transponder

GPS Receiver

Antennas

56

For the minimum cost model, the less expensive linear transponders would be the

best choice. Since this transponder is custom-designed, its cost would vary according to

the final design requirements. However, the best working design that would readily fit

into CUBIK is still less expensive than an FM transponder.

In order to achieve maximum performance, the type of transponder chosen

depends on what is to be maximized. For example, if the goal was to maximize the

amount of data transfer relative to power consumption, then an FM transponder would

provide maximum performance. Alternatively, if the primary concern is simultaneous

support for multiple users interfacing with the satellite, a linear transponder provides

maximum performance.

In order to achieve maximum simplicity, an FM transponder is used because their

chips are readily available for purchase and simpler to use than linear transponder chips.

4.4.2 Antennas This section presents the choice of antenna for each of the three design models.

When trying to minimize overall design cost, a single-plane antenna is chosen. As shown

in Table 4.6, single-plane antennas are less expensive. For example, the most expensive

single-plane antennas would generally cost one-third the price of a circular antenna.

Table 4.6 Antenna Characteristics

Antenna Type

Size

Deployment Difficulty

Cost

Percentage of Data Transfer

Number Needed

Circular

2m or 70 cm

Very difficult $300-$400

70% 80%

1

Single plane

0.5m - 1m

Simple $50- $100

50% 65%

3 or 4

57

For the maximum performance model, the circular antenna provides higher a

quality of reception. Also, circular antennas are able to send and receive signals with a

higher percentage of data transfer.

In order to achieve maximum simplicity, a single- plane antenna is chosen.

Single-plane antennas are smaller in size compared to circular antennas. Despite needing

more single-plane antennas, the circular antenna requires a sophisticated deployment

mechanism and would require more space.

4.4.3 GPS Receiver

This section discusses the advantages and disadvantages of different GPS

receivers. The GPS receivers being considered are listed in Table 4.7.

For the minimum cost design concept, the choice of a GPS receiver is not a major

factor since GPS receivers are similarly priced. For example, the RGM-2000 and RGM-

2001 are $500 each.

To maximize the performance we would chose RGM-3000 since it uses 20%

lower power and has 30% higher sensitivity to incoming signals than the other two.

RGM-3000 is capable of regulation of its own power efficiently.

For the maximum simplicity design model, the RGM-3000 would be the best

choice since it has an integrated path antenna. Built in path antennas allow operation

without an external antenna for the GPS receiver.

58

Table 4.7 Communications Summary

4.5 Power

The power subsystem must efficiently generate, store, regulate, and distribute

power to the other subsystems. The power subsystem must also supply a constant power

through eclipse, and provide enough power to meet the average and peak loads.

4.5.1 Power Generation

There are four types of photovoltaic cells being considered: silicon, indium

phosphide, gallium arsenide, and multi-junction gallium arsenide. Silicon is the least

efficient with an efficiency rating of about 14%. Indium phosphide is the second least

efficient with a rating of about 18%. Gallium arsenide is the next most efficient with a

rating of 19%.Multi-junction gallium arsenide cells are the most efficient with a rating of

22% or more. The cost of the cells increases with the efficiency rating.

For the minimum cost model, inexpensive silicon photovoltaic cells are used to

generate power. The maximum performance design model uses high efficiency multi-

junction gallium arsenide solar cells to generate power. Since the complexity of the solar

cells is similar, the maximum simplicity design will use high efficiency multi-junction

gallium arsenide solar cells.

Component Minimum Cost Maximum Performance

Maximum Simplicity

Transponder Linear

Linear FM

Antenna Single plane Circular

Single plane

GPS Receiver REB-2100 & REB-2000

RGM-3000 RGM-3000

59

4.5.2 Energy Storage Three different battery types are being considered for energy storage aboard

CUBIK: nickel cadmium, nickel metal hydride, and lithium-ion. As shown in Figure 4.3,

lithium-ion batteries are the most efficient of the three batteries. However, lithium-ion

batteries are also the most expensive. The nickel cadmium batteries are the cheapest, and

the nickel metal hydrides are priced between the lithium-ion and the nickel cadmium

batteries.

Figure 4.3 The performance characteristics of the battery types16. For the minimum cost model, the power is stored in inexpensive, widely available

nickel cadmium batteries. For the high performance and maximum simplicity models, the

more expensive but more efficient lithium-ion batteries are the choice for energy storage.

4.6 Attitude Determination and Control Systems

Attitude determination and control systems are directly influenced by the CUBIK

objectives and the needs of the customers. The CUBIK kit’s marketability is directly

affected by the form of attitude stability included with the kit. Marketability is increased

60

as the number of stable axes on the satellite is increased. Some forms of attitude

determination and control are not considered in the three design models.

4.6.1 Attitude Determination

Mass, size and cost of the various forms of attitude determination sensors

discussed in Chapter 3 are not known. Sensors for small satellites are not readily

available. Space tested components small enough to operate in the satellite are difficult to

find, which has led to an investigation of other potential sources of ADCS. Hobbyist

components used for control and determination of various types of remote controlled

models are being considered. Currently, the team is assuming suitable sensors are

available. The three CUBIK design models, coupled with this assumption, result in the

following three attitude determination models.

The request-for-proposal requires a sun-sensor only, therefore the minimum cost

objective design would just contain a sun-sensor. For initial determination, the sun-

sensor alone can determine the attitude of the satellite with relation to the Sun, but not the

Earth. Further attitude determination will need to come from another system, such as

communication. Further cost reduction is obtainable by using the solar panels on the

satellite to form the sun-sensor. An algorithm can determine the direction of the sun by

analyzing the current created by the varying strengths of radiation on each face.

The maximum performance model incorporates the RFP-required sun-sensor

with one or more other sensors to give the greatest determination accuracy. Combining an

industry constructed sun-sensor and a horizon sensor will give pitch, roll and yaw

reference data with a potential accuracy of +/- 0.01o.15 Cost has yet to be determined, but

estimation from similar instruments for much large satellites indicates a possible cost of

$1000+.

61

The simplicity model design is similar to the minimum cost design with the

exception that a professional sun-sensor is used instead of employing the solar cells. This

configuration does not require reconfiguring the determination algorithm if the solar cells

are modified by the customers. Potential accuracy is on the order of +/- 0.1o. Cost is

estimated from similar instruments to a possible cost of $500+.

The investigation of sensors is continuing and these designs may be modified at

any point in the future.

4.6.2 Attitude Control

Attitude control actuators for small satellites are difficult to find. Currently, the

CUBIK team is awaiting information from Ikarus hobby gyroscope manufacturer about

their product in regards to its possible use in space. Other possible sources are currently

being investigated.

Omitting the attitude control function is required to achieve the minimum cost

design. The omission of attitude control is allowed in the RFP, but limits the number of

possible payloads CUBIK can carry. Not having an attitude control capability also

impacts communications utilization of directional control to achieve minimum power use

when broadcasting to ground stations. The result of no attitude control on

communications is the possibility of the satellite tumbling in a manner that will prevent

effective communication. The absence of attitude control will result in a pointing

accuracy of +/- 180o on all axes. The lack of a control system can negatively impact the

communications and the mission.

The maximum performance model consists of a design for maximum pointing

accuracy. Three orthogonal momentum wheels give a pointing accuracy from +/- 1o to

+/- 0.001, depending on manufacturing.15

62

The maximum simplicity model is identical to the minimum cost model. The

three design models are tabulated in Table 4.8. The estimations in Table 4.8 are provided

as a possible, but not definitive, guideline to determining the final design..

Table 4.8 Relative Weights for ADCS Determination of the Three Models

Model Accuracy Cost Mission Effectiveness

Minimum Cost 1 9 3

Maximum Performance

9 1 9

Maximum Simplicity 3 7 3

4.7 Documentation

The documentation of the CUBIK specifications and directions for operation

should be clear, concise, and easy to understand. There are a variety of methods for

documenting the specifications and instructions. The different alternatives are analyzed

with respect to the three design concepts.

4.7.1 Documentation for TransOrbital The documentation for TransOrbital needs to include all the specifications so that

the company can readily produce the subsystems for sale. An inexpensive method for

documentation is to include the specifications and assembly drawings on a CDROM in

PDF format. Recordable CDROMs are inexpensive and documents can be converted to

PDF format at no expense. In order to maximize the performance of the documentation,

CAD files would be included. A simple method for documentation would also include

providing TransOrbital with a CD. The assembly drawings could be easily transferred

63

from the CD to a computer if TransOrbital needed to alter them or use them in any other

way.

4.7.2 Documentation for the User

Customers of CUBIK need the operating manual to be universal and easy to read.

A minimum cost instruction manual could be included on a CDROM in PDF format. As

stated before, CDROMs are inexpensive and provide an easy way to view the manual. A

video could be included with the manual to maximize the performance of the instructions.

An instructional video would cost more, but it would include clips which show the user

how to install the payload in the kit and how to remove any unnecessary components. A

simple method for documenting the instructions is to provide the users with a paper copy

of the operating manual. The manual allows for easy access because a computer would

not be necessary.

There are many choices for documentation of the CUBIK specifications and

operating procedures. Table 4.9 illustrates the best choices for each design model. The

table is a way to analyze how each choice would affect the CUBIK objectives.

Table 4.9 Documentation Choice for Each Design Model Documentation Design Concept 1 Design Concept 2 Design Concept 3 Documentation for TransOrbital

Specifications in PDF format on CD

Include CAD files with CD

PDF format on CD

Documentation for Customers

Manual in PDF format on CD

Operating video included with CD

Paper copy of manual

64

4.8 Summary System analysis presents three design concepts: a minimum cost model, a

maximum performance model, and a maximum simplicity model. The analysis is easier

to understand when specific models are made from the alternative solutions. These design

models are presented here in Table 4.10. The minimum cost model is designed to be the

least expensive possible working model. The maximum performance model has the

highest level of performance regardless of cost. Maximum simplicity models are

designed to have the simplest possible subsystems and subsystem interactions.

65

Table 4.10 Summary of the Three CUBIK Design Models Subsystem Minimum

Cost Model

Maximum Performance Model

Maximum Simplicity Model

Structure

Material Aluminum Aluminum Aluminum Panels Solid sheet Perforated/Framework Solid sheet Rails/Posts Extruded solid Square tube L-shape Internal Structure PC104 Custom-designed PC104

Flight Computer Microcontroller Microprocessor Microcontroller

Flight Computer

Software

Operating System uClinux

Linux

uClinux

Languages C++ C++ C++

Communications

Transponders Linear FM Linear Antennas Single Plane Circular Single Plane GPS Receiver REB-2000 & REB-

2100 RGM-3000 RGM-3000

Power

Power Generation Silicon High efficiency multi-junction Gallium Arsenide

High efficiency multi-junction Gallium Arsenide

Power Storage Nickel Cadmium Lithium Ion Lithium Ion

ADCS

Sensors Solar-cell sun sensor & communication system

Sun-sensor, horizon sensor

Sun-sensor

Control No control 3 momentum wheels No control

Documentation

For TransOrbital PDF file on CD CD include CAD designs

PDF file on CD

For Users PDF format on CD Operating video included with CD

Paper copy of manual

66

Chapter 5: The Future of CUBIK

5.1 Semester Summary

CUBIK is a picosatellite which is designed to be low cost, lightweight, and

operate with a variety of customer chosen payloads. The payload for CUBIK can be

anything from a simple science experiment run by high school students learning about

space to a more complicated device designed by professional scientists or engineers.

Chapter 1 gives a history of satellites and describes the scope of CUBIK. The chapter

also identifies the needs, alterables, and constraints for the project. Chapter 2 outlines

CUBIK’s objectives and establishes measures of effectiveness to evaluate the objectives.

The main objective is to design a universal affordable CubeSat bus kit, where the three

top-level objectives under the main objective are: maximize performance, minimize cost,

and maximize simplicity. Chapter 3 reviews the six subsystems of CUBIK and describes

alternative methods for achieving the objectives of each subsystem. Chapter 3 also

discusses the feasibility of the alternatives in relations to the needs and constraints.

Chapter 4 developed three design concepts for CUBIK. These concepts were used as

models to evaluate the alternative solutions developed in Chapter 3. The information

contained in these chapters will be used necessary for further development of the CUBIK

design.

67

5.2 Future Plans

5.2.1 Structure The next stage of development for the structure subsystem will be a review of the

structural component requirements presented in the problem definition. Following this

review, the alternatives selected for each design model will be re-evaluated to determine

if the all the design objectives are met. Considerations will include the impact of the

structural alternatives on other subsystems.

Following the re-evaluation, research will be conducted to determine the

availability and cost of structure components. This research will include manufacturing

costs associated with producing components, which are not commercially available. If

resources are available, arrangements will be made to purchase the materials necessary to

build the prototype CUBIK structure.

Computer aided design (CAD) of the structure will be the next stage of structure

development. The alternative design models will be tested using computer simulation

programs such as Unigraphics and ADAMS20, 21. Testing will determine theoretical

structural performance under expected launch and operating conditions.

Testing the structural components and their interactions with other subsystems

will require the fabrication of design prototypes. Initial prototypes will be simple forms

which facilitate the design of the other subsystems and allow physically testing design

alternatives. The structural design development culminates in the fabrication of a final,

“working” structural prototype.

5.2.2 Flight Computer

68

This semester the constraints and the alternatives for the flight computer were

reviewed. The various solutions for the problem at hand were discussed and examined.

Next semester the flight computer subsystem will be extended with decisions being made

about the kind of processor to be used, the kind of memory needed and the entire

computer architecture for the CubeSAT will be laid out. The alternate solutions will also

be revisited with the aim of excluding redundant solutions and including the alternatives

that are more likely to be used in the unit. Compatibility of the components will be a

major factor in deciding the parts to use.

We aim at building a working prototype for the CubeSAT by the end of the next

semester. Hence, after all the components are selected, they will be put together to form

the complete BUS for the CubeSAT. It is this period when the Operating system that will

drive the CubeSAT will be written and tested for proper functionality

5.2.3 Flight Computer Software

Up to this point we have defined the requirements and specifications of each

subsystem of the CubeSat. From these requirements we were then able to identify

several different possible alternatives for the design that would fit the constraints. These

alternatives will be studied in greater detail in the next course of research which will lead

to the final design of the CubeSat.

In the particular case of the FCS, we now know the tasks that the FCS will have to

carry out and we already have alternatives that handle these tasks. From these

alternatives, others will be added in the next research phase in order to have a greater

panel of possibilities to choose from. One of these new possibilities will be for us to

design our own RTOS. However, the phase of finding the alternatives will be short so

69

that we could rapidly determine the RTOS for the final design and focused ourselves on

the development of the software and their installation in the spacecraft flight computer.

5.2.4 Communications The next stage of development for the communications subsystem will be a

complete re-evaluation of each component in the system with respect to the problem

definition. Complete re-evaluation will be performed to each of the alternatives selected

to determine if they fall within the design objectives. A number of detailed

considerations will also be made regarding how the communication subsystem will affect

the other subsystems.

Once finished with the re-evaluation stage we will carry out an extensive research

on communication components to determine the best options from a point of view that

meets our requirements. The research will take into consideration cost, availability,

functionality, simplicity (in terms of use), and volume of the communications

components. The consideration priorities will be discussed in details. The choice of the

alternative and the final component will be determined with regard to the interactions it

has with the other subsystem components.

Various computer-aided software’s will be used to generate detailed equations

and circuits modeling the components chosen. Various theoretical tests will be

performed on the component regarding voltage and current capabilities. Interactions with

other systems might be modeled also on a theoretical basis.

After decisions considering which components CUBIK will be considering,

physical testing of the components will be performed. Different tests like communication

range, power requirements, heating, radiation effects and others will be performed.

70

5.2.5 Power

Next semester the power subsystem will be going through a second iteration of the

design process, a prototyping phase, and then an integration phase. The second iteration

of the design process will decide which components are viable alternatives, and which

components appear to be most feasible. Costs and availability will be determined for

each alternative. The alternatives will be examined in relation to the components being

considered for the other subsystems and a final choice will be made on which alternatives

will be used. This process is planned to be completed within five to seven weeks.

Once specific components are chosen, the power subsystem will go through a

prototyping phase. During this phase the power subsystem will actually be designed and

prototyped. The prototyping phase is expected to take eight to ten weeks. During this

phase all bugs and compatibility issues must be worked out. The prototyping phase will

be limited by cost unless funding is obtained. Once a working prototype of the power

subsystem is finished, the process of integrating it into the satellite with the other

subsystems will begin.

5.2.6 Attitude Determination and Control Systems The next stage of development for the ADCS subsystems will begin with a re-

evaluation of ADCS in the problem definition. The re-evaluation will take into

consideration all that was determined since the first evaluation of the problem definition.

Considerations will include the impact of ADCS on other subsystems, cost, and on the

marketability of the CUBIK kit.

Following the re-evaluation, there will be an extensive search for currently

developed ADCS systems. The search will focus on systems that can meet the mass, cost

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and volume constraints of CubeSat. The CubeSat constraints may make it necessary to

consider the development of ADCS systems built especially for CubeSat and the pico-

satellite industry. When all available systems have been determined, the different options

will be evaluated in terms of the constraints, subsystem interactions and objectives. The

result of the evaluation will determine the final configuration for the ADCS system for

the CUBIK kit.

Control programming for the ADCS final configuration will be developed using

stability and control algorithms. Variables for the algorithms will be determined by

analysis. The analysis and programming will be developed iteratively.

The final configuration will be analyzed in terms of dynamic reactions to stimulus

such as solar winds, eclipse effects and structural vibration caused by thermal and

radiation effects. The analyses may include computer simulation, electronic simulation

and physical prototype testing. Pointing error and transient response time will be the

primary measures of effectiveness.

A final design and prototype will be developed for presentation to TransOrbital if

there is sufficient time to do so before the end of the semester.

5.2.7 Documentation

Documentation of all specifications, diagrams, and directions for operation must

be recorded during the upcoming semester. Beginning with the optimization stage and

continuing into the decision stage, all data and specifications for components of each

subsystem must be recorded. These values will be included in the documentation to

TransOrbital and the customer.

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Once a final design is concluded upon, diagrams of each subsystem and their

internal components must be created. Diagrams can be created with a CAD program.

The specifications and diagrams will have to be organized into a clear and concise

manual which TransOrbital can utilize to produce the subsystems for sale. Late in the

semester, a choice will have to be made on what format to use for documentation for

TransOrbital.

After the prototype is constructed, a clear, concise, easy to understand instruction

manual for the customers will be made. A decision whether to produce the manual on

CDs or using paper copies will be made.

5.3 Summary

The next semester will be divided into phases. Each phase has been allotted an

estimated amount of time for completion. The first phase will be a second iteration of the

design process. Concurrent with that stage we will be determining cost and availability

of subsystem components. A subsystem prototyping phase will begin as specific

components are chosen. During this stage, complete subsystem prototypes will be built

and tested. After costs and availability have been determined, we will begin optimizing

and determining component compatibility. As complete subsystems are constructed, they

will be integrated into a working bus prototype. Appropriate documentation of

subsystem specifications and diagrams will be produced during the subsystem integration

phase. After a working prototype is constructed, a complete operating manual for the

customer will be written.

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