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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
ii
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
vii
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.
2
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.
5
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.
7
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
8
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
9
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
10
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.
11
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
12
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
13
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
14
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.
15
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 ($)
MOE
Minimize
Maximize
Legend
Price ($)
16
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.
17
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
18
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
19
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.
20
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
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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.
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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.
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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.
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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.
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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.
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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
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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.
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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.
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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
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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
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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.
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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
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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+.
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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
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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
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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
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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.
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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
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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.
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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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References 1 Isakowitz, S. J., Hopkins Jr., J. P., Hopkins, J. B., International Reference Guide To Space Launch Systems (Reston, VA: American Institute of Aeronautics and Astronautics, 1999). 2 Wertz, J. R., and W. J. Larson, ed., Reducing Space Mission Cost (Torrance, CA: Microcosm Press, 1996), pp.119-123. 3 Heidt, H., J. Puig-Suari, A. Moore, S. Nakasuka, R. J. Twiggs, "CubeSat: A New Generation of Picosatellite for Education and Industry Low-Cost Space Experimentation," 14th Annual USU Conference on Small Satellites, Logan, Utah, August 2000. 4 One Stop Satellite Solutions. “OSSS CubeSat Kit.” 16 July 2002. http://www.osss.com/products/cubesat.htm 5 Stanford University. “CubeSat Design & Test Requirements.” 14 March 2001 http://ssdl.stanford.edu/cubesat/design.html. 6 Foley, T., "Financial Turmoil & Shakeouts Come to the Satellite Sector in 2002," Via Satellite's 2002 Satellite Industry Directory (Potomac, MD: PBI Media, LLC, 2002), pp. 1-6. 7 Internal Amateur Radio Union and Radio Amateur Satellite Corporation. “Information for Prospective Owners and Operators of Satellites Utilizing Frequencies Allocated to the Amateur-Satellite Service.” 2002 http://www.iaru.org/satellite/prospective.html. 8 Tokyo Institute of Technology Lab for Space Systems. “Cubical Titech Engineering Satellite.” 2002 http://horse.mes.titech.ac.jp/srtlssp/cubesat/index.html. 9 LinuxDevices.com. "Linux on Microcontrollers." http://www.linuxdevices.com/links/LK8053710489.html 10 Beer, F. P., Johnston, E. R. Jr., Mechanics of Materials (New York, NY: McGraw-Hill, Inc, 1992), pp. 700-703. 11 Piscane, V. L., Moore, R. C., Fundamentals of Space Systems (NewYork, NY: Oxford University Press, 1994), pp. 504-514. 12 Principal Metals. “Property Database.” October 2002 http://www.principalmetals.com/properties/step1.asp 13 PC104.org, “PC/104 Specification Version 2.4. PDF.” 29 May 2002. http://www.pc104.org/technology/PDF/PC104Specv246.pdf
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14 Incropera, F. P., DeWitt, D. P., Heat and Mass Transfer (New York, NY: John Wiley & Sons, Inc., 2002), pp. 700-756. 15 Wertz, J., Larson, W., Space Mission Analysis and Design, Third Ed. (El Segundo, CA: Microcosm Press and Kluwer Academic Publishers, 1999) 16 Toshiba-Europe. “Toshiba Visions: Power in a Pint Pot,” 15 September 2001. http://www.toshiba-europe.com/computers/sna/tnt/visions96/power.htm 17 Microsoft Corporation. “Windows CE .Net Home.” 14 November 2002. http://www.microsoft.com/windows/embedded/ce.net/ 18 Intel Corporation. “Technical Specifications.” November 2002. http://www.intel.com/intel/intelis/museum/exhibit/hist_micro/hof/tspecs.htm 19 Principal Metals. “Magnesium AZ31B.” 2002. http://www.principalmetals.com/properties/result.asp?Family=Magnesium+Alloy s&MetalName=AZ31B 20 EDS. “EDS Products – Unigraphics.” 2002. http://www.eds.com/products/plm/unigraphics/ 21 MSC Software. “MSC.ADAMS Software Homepage.” 2002 http://www.adams.com/
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