The document proposes a CubeSat design called SPDR Sat to help address the growing issue of space debris in low Earth orbit. SPDR Sat would be a 3U+ CubeSat that uses commercial off-the-shelf hardware, including a proprietary net as its capture device. It is designed to complete orbital maneuvers to rendezvous with and deorbit a single piece of space debris approximately 9000cm3 in size. The seven subsystems addressed are thermal management, power, communication, computing, propulsion, attitude determination and control, and structure. The goal is to provide a low-cost solution to space debris removal and help delay the potential Kessler Syndrome effect.

1. Transatmospheric Vehicle Design
Professor Kurt Anderson
Rensselaer Polytechnic Institute
Space Debris Removal Satellite
Group 5
Alex Link: Thermal Management
Andrew Rapsomanikis: Structural & Deployable Systems
Greg Black: Propulsion System
James Leith: Communications System
Jeff Lehrer: Primary Computing System
Jordan Fisher: Attitude Determination & Control System
Shawn Madden: Power System

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Executive Summary
There currently exists a concerning amount of debris residing in Low Earth Orbit (LEO).
As time goes on, the space debris will continue to build up, and collisions will continue to
become increasingly likely, decreasing the prospect of future missions in the region.
Intervention is required, and the solution comes most efficiently in the form of a CubeSat.
NASA’s Jet Propulsion Laboratory (JPL) and Rensselaer’s Aerospace Systems Solutions, Inc.
(RASSI) have identified this issue and are seeking a cost effective solution.
The design of a CubeSat ultimately depends on its mission requirements. Due to a
CubeSat’s low cost, size, and mass it is the prime candidate for efficient removal of space debris.
However, the environment of LEO is unforgiving and the CubeSat must be able to operate in a
safe temperature range and provide enough power to all of its components throughout the
mission. If the CubeSat can survive and perform the required maneuvers to complete the
mission, the issue of space debris can be resolved in the coming decades.
Space Debris Removal Satellite (SPDR Sat) is the proposed solution. The SPDR Sat is
designed to deorbit a single piece of space debris sized approximately 9000cm3
in size. The
design utilizes primarily Commercial Off-The-Shelf (COTS) hardware with the addition of a
proprietary net, serving as the designated capture device. The utilization of COTS hardware
provides a cheaper alternative to proprietary hardware, as well as a way to easily replicate the
final design. The development effort emphasized structural design, thermal and power
management, propulsion, attitude control, communication and computing systems and
fabrication of the net device.
Through the use of Systems Tool Kit, STK, Solidworks and other modeling/simulation
tools, the SPDR Sat team has proved that their design is capable of completing the necessary
orbital maneuvers, adjusting its orientation to face the target debris for capture, managing and
powering each of the individual components, operating in a safe temperature range and
maintaining communication with a set of different potential ground stations.

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Table of Contents
Executive Summary 2
List of Figures 5
List of Tables 7
Terms and Abbreviations 8
1 Introduction 9
1.1 CubeSat History 9
1.2 Problems to be Addressed 10
1.3 Justification & Benefits 10
2 Subsystem Division 12
3 Subsystem Requirements & Performance Specifications 13
3.0.1 Mission Walkthrough 13
3.0.2 Mission Success 14
3.1 Structure and Deployables 14
3.2 Power Requirements 15
3.3 Propulsion Requirements 16
3.4 Attitude Determination and Control System Requirements 16
3.5 Primary Computing Requirements 17
3.6 Communication Requirements 17
3.7 Thermal Requirements 17
4 Subsystem Concept Development 20
4.1 Structural Concept Development 20
4.2 Power Concept Development 21
4.2.1 Battery Selection 21
4.2.2 Solar Panel Selection 23
4.3 Propulsion Concept Development 24
4.4 Attitude Determination and Control System Concept Development 25
4.5 Primary Computing Concept Development 27
4.6 Communications System Concept Development 29
4.7 Thermal Concept Development 32
5 Design Analysis 36
5.1 Structures System Analysis 36
5.2 Power Systems Analysis 37
5.3 Propulsion Systems Analysis 40
5.3.1 Hohmann Transfer 40
5.3.2 Phasing Maneuver 42
5.3.3 Inclination Change 46
5.3.4 Desaturation of Momentum Wheels 48
5.4 ADCS Systems Analysis 49
5.4.1 Stability and Control 49
5.4.2 Pointing Accuracy 53

4. 4
5.4.3 Sun Tracking 54
5.5 Primary Computing Systems Analysis 55
5.6 Communications Systems Analysis 56
5.7 Thermal Systems Analysis 59
6 Final System Design Overview 64
6.1 Structures System Overview 64
6.2 Power System Overview 69
6.3 Propulsion System Overview 69
6.4 ADCS System Overview 69
6.5 Primary Computing System Overview 70
6.6 Communication System Overview 72
6.7 Thermal System Overview 73
7 Cost Analysis 75
8 Risk Analysis 76
9 Future Problems 78
10 Future Applications 79
11 References 80
Appendix A: Propulsion Analysis Code 84
Appendix B: Thermal Analysis 90
Appendix C: ADCS 92
Appendix D: Communications 99
Appendix E: Power Systems 101
Appendix F: Structure 105
Appendix G: Component Spec Sheets 108
Contributions 115

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List of Figures
Figure 1.1.1: ESA Fishing Net Concept
Figure 2.1: Subsystems Setup
Figure 3.1.1: 3U+ P-Pod “Tuna Can” Allotted Space
Figure 4.1.1: RadiusSpace Frame
Figure 4.4.1: Cube ADCS
Figure 4.6.1: NanoCom U482C UHF Half-duplex Transceiver
Figure 4.6.2: ISIS VHF downlink / UHF uplink Full Duplex Transceiver
Figure 4.6.3: ISIS UHF downlink / VHF uplink Full Duplex Transceiver
Figure 4.6.4: ISIS Deployable Antenna System
Figure 4.7.1: Conduction between two surfaces
Figure 4.7.2: Liquid cold plate heat sink
Figure 4.7.3: Structure of MLI blanket including common materials in each layer
Figure 5.2.1: Lighting Times at 400 km Orbit
Figure 5.3.1: Propellant Consumption vs Radius for Hohmann Transfer
Figure 5.3.2: Delta-V Requirement vs Radius for Hohmann Transfer
Figure 5.3.3: Propellant Consumed vs Revolutions in phasing orbit
Figure 5.3.4: Delta-V requirement vs Revolutions in phasing orbit
Figure 5.3.5: Propulsion used vs. Inclination change
Figure 5.3.6: Delta-V requirement vs Inclination change
Figure 5.3.7: Matlab Output
Figure 5.4.1: Angular Velocity vs. Time Stowed Solar Panels
Figure 5.4.2: Euler Parameters vs. Time Stowed Solar Panels
Figure 5.4.3: Angular Velocity vs. Time Deployed Solar Panels
Figure 5.4.4: Euler Parameters vs. Time Deployed Solar Panels
Figure 5.4.5: Solar Angle
Figure 5.5.1: Remote image processing tests
Figure 5.6.1: Kauai Community College Ground Station
Figure 5.6.2: Access Times on Jul 4 2017
Figure 5.6.3: Visual Representation of Access Times on Jul 4 2017
Figure 6.1.1: Expanded View of SPDR Sat
Figure 6.1.2: SPDR Sat Stage 1
Figure 6.1.3: SPDR Sat Stage 2
Figure 6.1.4: SPDR Sat Stage 3
Figure 6.1.5: SPDR Sat Stage 4
Figure 6.4.1: Cube ADCS
Figure 6.5.1: CubeComputer
Figure 6.5.2: CubeSense

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Figure 6.5.3: ESTCube-1 Camera
Figure 6.5.4: Control Flowchart
Figure 6.6.1: SPDR Sat trajectory when passing over KCC over a day
Figure 6.6.2: Broadcast signal from SPDR Sat

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List of Tables
Table 3.7.1: Minimum and maximum allowable temperatures for operation
Table 4.2.1: Battery Specifications Sheet
Table 4.2.2: Battery Selection Matrix
Table 4.2.3: Solar Panel Specifications Sheet
Table 4.2.4: Solar Panel Selection Matrix
Table 4.3.1: Propulsion Selection Matrix
Table 4.3.2: Properties of hydrazine and FLP-106
Table 4.4.1: ACDS Performance Specifications
Table 4.4.2 ADCS Selection Matrix
Table 4.5.1: CPU Considerations
Table 4.5.2: CPU Selection Matrix
Table 4.6.1: Concept Selection Matrix for Transceiver
Table 4.7.1: Thermal Component Selection Matrix
Table 5.1.1: Part List for Mass and Volume Analysis
Table 5.2.1: Various Power Setups
Table 5.2.2: Total Power Consumption Average
Table 5.3.1: Hohmann Transfer Propulsion Analysis
Table 5.3.2: Propulsion Analysis for Phasing Maneuver
Table 5.3.3: Propulsion Analysis for Inclination Change
Table 5.3.4: Momentum Arm Thruster Force
Table 5.4.1: Inertia Matrix for Stowed Solar Panels
Table 5.4.2: Detumble Parameters
Table 5.4.3: Inertia Matrix for Solar Panels Deployed
Table 5.4.4: Control Simulation Setup
Table 5.4.5: Solar Panel Efficiency as a Function of Angle
Table 5.5.1: Component Interface
Table 6.7.1: MLI Materials and thicknesses
Table 7.1: Cost Analysis
Table 8.1: Risk Severity and Probability

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Terms and Abbreviations
 ADCS – Attitude Determination and Control System
 AEOLDOS – Aerodynamic End of Life De-Orbit System
 CAD – Computer Aided Design
 COTS – Commercial Off-The-Shelf
 CPU- Central Processing Unit
 ESA – European Space Agency
 FEA – Finite Element Analysis
 GEVS – General Environmental Verification Specification
 IEEE – Institute of Electrical and Electronics Engineers
 ISIS – Innovative Solutions in Space
 ISS – International Space Station
 ITU – International Telecommunication Union
 JPL – Jet Propulsion Laboratory
 KCC- Kauai Community College
 LEO – Low Earth Orbit
 MATLAB- Matrix Laboratory
 MLI – Multi-Layer Insulation
 P-POD – Poly Picosat Orbital Deployer
 RASSI – Rensselaer’s Aerospace Systems Solutions, Inc.
 SPDR Sat – Space Debris Removal Satellite
 STK- Systems Tool Kit
 UHF – Ultra High Frequency
 VHF – Very High Frequency

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1 Introduction
1.1 CubeSatHistory
Space debris has been an ongoing problem for many years now. Jettisoned rocket stages,
decommissioned satellites and random debris particles have been collecting in various orbits
around Earth. A theory developed by Donald J. Kessler, known as the Kessler Syndrome, states
that there exists a critical density of objects in LEO beyond which a cascading collision effect
could occur, wiping out local spacecraft and rendering the region unviable for occupation for
generations (Innocenti 7). As of 2014, there were 300,000 pieces of tracked debris in LEO (ESA
About Space Debris). In other words, the critical density of which Kessler spoke is being
approached and without proper discretion his hypothetical scenario could come to realization.
This is something that must be avoided at all costs. A few companies/research groups have
come up with a few proposed solutions to help combat the problem of space debris. A group at
Texas A&M has designed a satellite called the “Space Sweeper with Sling-Sat”. This satellite
would harness the momentum imparted by capturing and ejecting one body to sling-shot to the
net piece of space-junk (Space-Junk). Swiss Space Systems has come up with a satellite known
as the Clean Space One satellite. This is a nanosatellite that uses a claw to grab onto space junk
and subsequently deorbiting to burn up in the atmosphere (Coppinger). The European Space
Agency has also developed an expanding foam to be sprayed on space debris, increasing the
surface area along with the coefficient of drag, allowing the space debris to fall into the Earth’s
atmosphere and burn up (Andrenucci).
The SPDR sat design is split into seven subsystems. The first subsystem is thermal
management. The satellite will need to keep all of its components within their functional range of
temperature. The second subsystem is power. The electric energy storage and voltage must allow
for all components to work while also allowing the solar panels to recharge. The third and fourth
subsystems are communication and computing. The satellite must be able to receive data and act
upon the information given from the ground. The fifth subsystem is propulsion. The propulsion
system must be able to control the satellite and move in space to a piece of space debris. This
goes along with the sixth subsystem, controls and attitude dynamics. This subsystem consists of
the algorithms and code that will keep the satellite oriented and allow a rendezvous with a piece
of debris. The final subsystem is structure and deployables. This consists of the frame as well as
how all of the other subsystem components will fit in the CubeSat.
The CubeSat will use commercial off the shelf hardware for all components except the
capture device. The capture device will resemble the European Space Agency “fishing net”
design, but for a smaller satellite and smaller pieces of debris. Due to the European Space
Agency’s (ESA) success on this design, which has been tested in microgravity parabolic-arc test
planes, a similar success can be expected for the design with only a reduction in scale. The
concept for the net design can be seen in the images below of the ESA “fishing net”.

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Figure 1.1.1: ESA Fishing Net Concept
1.2 Problems to be Addressed
As part of the original requirements of this project, the team designed a satellite which was
comprised predominantly of COTS hardware. The goal of this is to ensure that all parts are
flight-tested, and ideally flight-proven. The overall goal of the design is to find a low-cost and
mass-producible method of lowering the amount of dangerous space debris in Low Earth Orbit.
Shown below is a list of the detailed requirements which the SPDR satellite will address.
 Conform to the CubeSat Standards
 Meet size requirements of either 3U+ CubeSat
 Endure a Poly Picosat Orbital Deployer (P-Pod) deployment with no damage sustained to
vital systems
 Insure no safety concerns during or after launch from cargo trunk
 ACS system able to stabilize and de-tumble satellite after deployment
 Ability to relay sufficient amounts of data during communication windows with ground
station
 Carry sufficient fuel for orbital maneuvers such as inclination changes, Hohmann
transfers and phasing maneuvers
 ACS system able to orient satellite for sun tracking and orbital maneuvers
 Solar panels and batteries able to provide enough power to run satellite systems such as
attitude determination and control, communication, and propulsion
 Thermal control system must enable the satellite to endure the extreme high and low
temperatures associated with its orbit
 After transfer maneuvers, satellite will be able to autonomously locate, track, and
approach target debris
 Proprietary capture device is able to secure the target and hold on to it
 Drag sail is able to decrease the satellites de-orbit time to less than 5 years
1.3 Justification& Benefits
As mentioned above, the Kessler syndrome predicts a cascade of collisions once Low
Earth Orbit reaches a critical density of debris. The exact value of this critical density remains
unknown, but if no action is taken within the near future, the Kessler syndrome may become
unavoidable. The increased collision probability would pose a larger threat to current and future,
satellites and spacecraft missions.
This issue could be addressed one of two ways: design all future space vehicles with
sophisticated collision avoidance systems or design a method to manage space debris.
Unfortunately, once the density of space debris in LEO surpasses its critical level, the collision

11. 11
avoidance systems will quickly become overwhelmed and rendered useless. These systems are
overly complicated and fail to address the root cause of the issue. On the other hand, designing a
method to deorbit space debris is a better solution which focuses on the long term benefit.
The proposed design for space debris removal, in this report, would provide several
benefits to NASA’s Jet Propulsion Laboratory. Decongesting LEO would not only help delay the
onset of Kessler syndrome, but it would also free up orbits for new satellites and spacecraft.
Cost, as with any space mission, is one of the largest design constraints. Using CubeSats saves
the client money on material costs and launch costs since these satellites are small relative to
most spacecraft and can be launched along with other missions. These spacecraft are also easily
mass-producible since they are composed of primarily COTS hardware. Furthermore, CubeSats
are unmanned which means human lives would not be risked in the process of debris removal.

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2 Subsystem Division
The SPDR Sat design is split into seven subsystems.
Figure 2.1: Subsystems Setup
The first subsystem is thermal management. The satellite will need to keep all of its components
within their functional range of temperature. The second subsystem is power management. The
electric energy storage and voltage must allow for all components to work while also allowing
the solar panels to recharge. The third and fourth subsystems are communication and computing.
The satellite must be able to receive data and act upon the information given from the ground.
The fifth subsystem is propulsion. The propulsion system must be able to control the satellite and
move in space to a piece of space debris. This goes along with the sixth subsystem, controls and
attitude dynamics. This subsystem consists of the algorithms and code that will keep the satellite
oriented and allow a rendezvous with a piece of debris. The final subsystem is structure and
deployables. This consists of the frame, capture device, and drag sail, as well as how all of the
other subsystem components will fit in the CubeSat and follow the regulations. Each of these
subsystems was chosen by team members based on their strengths and preferences.

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3 Subsystem Requirements and Performance Specifications
3.0.1 MissionWalkthrough
In the journey from launch to nominal performance to mission completion, the design of
the CubeSat is to allow for different components to be turned on and off. The required
components in use are determined not only by the mission requirements of the CubeSat, but also
the power requirements needed to run all the onboard systems. In order to fully understand the
components to be chosen, a step-through of the different aspects of the detailed mission are as
follows.
 Launch from P-Pod
o CubeSat will need to detumble from an unknown orientation and spin. The
detumbling will be accomplished using a combination of the onboard gyros and
the reaction wheels to give a stead orientation.
o Solar panels and antenna will delay deployment until stable orientation is found
o Once stable orientation is achieved, solar panels will extend and lock into place
followed by antenna deployment
 Pre-Nominal Orientation
o Satellite will find and orient itself with its solar panels facing directly the sun
using sun sensor (located on same face as panels)
 Control loop will run to find optimum 90 degree orientation to sun
providing max power to systems and recharge any battery loss
o Once communication is established with ground station, satellite will attempt to
orient itself in space using sun and earth sensors in conjunction with ACS
o System will run diagnostics and downlink any data of errors to ground station
 Nominal Operation
o Depending upon true power requirements, ACS will be activated on an interval
(such as every 30 sec) to reorient panels towards sun
 Will allow for efficient use of battery ensuring minimal degradation of
battery life-cycle
 When within umbra, reorientation will be in one or two steps in
preparation of sunlight
 If batteries are at minimal drain, ACS will be suspended and sun tracking
will not occur until necessary
o Uplink and Downlink will occur once or twice per day to ensure nominal
performance and allow for any optimized code or update instructions
 Mission
o Once a target debris has been identified by ground station, rendezvous maneuver
instructions will be uplinked to satellite
o When instructions uplinked, satellite will ensure batteries are at full charge
o Maneuver execution through successive firings of propulsion followed by sun
tracking during non-firing times
o Autonomous in-close rendezvous performed including net firing and capture
 End Life/ End Mission
o Following confirmed capture, drag sail extended
o Solar panels will continue recharge of batteries

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 No sun tracking will be performed
o Intermittent updates will be performed if necessary/ possible through uplink and
downlink with ground station
Overall end of life rules are determined by NASA-STD 8719-14 RevA
3.0.2 MissionSuccess
For the SPDR Satellites mission to be considered a success, the following capabilities must be
demonstrated.
 Able to autonomously track and maneuver towards debris in local vicinity using image
processing.
 Net capture device able to entangle target and remain attached for the duration of the
mission
 Satellite able to deorbit with debris in less than five years after entanglement
 Satellite built primarily primarily of COTS hardware, is low-cost and mass-producible
3.1 Structure and Deployables
The structure of a CubeSat has a set of basic constraints. The first is on the size. The
CubeSat must fit within the P-Pod. This allows a 3U to be 10 centimeters in width and height
and 30 centimeters in length. The CubeSat must also weigh no more than 4 kilograms. The
center of mass is required to be within 2 cm center in width and height, and 7 centimeters in the
lengthwise direction from the geometric center. In Appendix F, the SPDR Sat can be seen to
follow the regulation in all three directions.
The SPDR Sat is considered a 3U+ CubeSat. This is because the propulsion system thrusters
protrude out of the 3U allotted volume. However, this is allowed because the P-Pod has a “tuna
can” of extra space within the spring area of the P-Pod. The four thrusters will protrude on the
inside of the compressed spring, but will not negatively affect the P-Pod launch. This allotted
3U+ volume can be seen below.

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Figure 3.1.1: 3U+ P-Pod “Tuna Can” Allotted Space (CubeSat Design Specification)
No pyrotechnics are allowed on board due to risk of explosion. A CubeSat cannot impair damage
to the must more expensive launch main mission. The SPDR Sat deployable components are all
deployed with the use of springs and electrical motors.
3.2 PowerRequirements
Power systems are the life blood of everything which travels into space. Without enough
power, a satellite in orbit or transit will have a complete mission failure. As such, a robust and
reliable enough power system must be found for any successful CubeSat design. The SPDR
satellite has been designed for a mission on a circular orbit of 600 km in altitude. To be
discussed later will be specifics about these orbit parameters but in general a single orbit will
occur around every 90 minutes. In these 90 minutes, the satellite will experience both full
sunlight as well as full eclipse from the Sun, Earth, and satellite movements. The amount of time
spent in sunlight and eclipse, or umbra, will affect the selection of the power storage and power
regeneration method.
While in eclipse, the storage product chosen, the battery, must have a substantial enough
capacity to effectively provide power for the various other subsystems which may or may not be
operating during eclipse. For the solar panels, they must be able to both continue to provide
power for these subsystems as well as recharge the depleted battery pack(s) in preparation for
another eclipse pass. Optimally, the solar panels will be able to fully recharge the battery during
the sunlight portion so that there will not be an overall drain on the battery leading to an eventual
complete depletion during normal mission operation.

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There were a few design constraints set upon the power system because of its mission
critical nature. The two main criterion focused on were the two discussed just above, that there
would be enough storage for umbra and enough solar panels to regenerate power in sunlight to
completely charge the depleted batteries. Meeting these two criterion was placed above all else.
But some other criteria looked at was form factor (size and weight), in space testing/ reliability,
and of course price. In the team’s selection of the system, there were much more robust systems
which could have been chosen. But these solutions were at a price point which would make the
overall cost of the satellite be unreasonable for the mission outcome (atmospheric burnup).
3.3 Propulsion Requirements
The propulsion system must be able to sufficiently maneuver the SPDR satellite throughout
Low-Earth-Orbit in order to get into range of specified critical space debris. The system must be
able to desaturate momentum wheels if needed. It must comply with CubeSat propulsion Design
Specifications which include the following (CubeSat Design Spec):
I. Any propulsion system shall be designed, integrated, and tested in accordance with
AFSPCMAN 91-710 Volume 3.
II. Propulsion Systems shall have at least 3 inhibits to activation
3.4 Attitude Determination and Control SystemRequirements
The system requirements for the Attitude Determination and Control system (ADCS)
come from the need to make sure the spacecraft can stabilize itself and meet the various pointing
accuracy requirements for different parts of the mission. Additional physical limitations were put
on the system to ensure that it was not too heavy or large. While designing the SPDR Sat weight
and size budgets were very tight so it was decided that the ADCS must not weigh more than 1kg
or take up more than 1U in size. Additionally, as to not take too much power, the ADCS was
limited to a maximum draw of ~5 watts. This size and power limitation represents the reaction
wheels, sensors, and any chips and Central Processing Units (CPU) required to control the craft.
The three performance requirements for the ADCS fell into the categories of torque
capabilities, pointing accuracy, and stability control. For torque control, the ADCS must be able
to produce enough torque to de-tumble the satellite while the solar panels are stowed, and also be
able to rotate the craft for sun-tracking and propulsive maneuvers when the solar panels are
deployed. More detailed analysis of the ADCS torque performance is covered later in the report.
Since an omnidirectional antenna will be employed on the SPDR Sat, no pointing accuracy is
required for communication. For the solar panels, a pointing accuracy of 5° will suffice, however
for orbital maneuvers and debris capture the pointing accuracy for the craft needs to be 1°.
Stability control is very important in that if the reaction wheels become saturated there must be a
fail-safe system in place aboard the satellite. There will be no room for dedicated ACS thrusters
aboard the craft, so the task of desaturating the reaction wheels will fall on magnetic torque tubes
built in to the ACS system.

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3.5 Primary Computing Requirements
Any component that it the “brain” of a system must be able to communicate and control
all parts of that system. For SPDR there are 3 main mission sections that the spacecraft will need
to be in control of. The first section is the attitude sensing and maneuvering. This includes the
ACS board, the momentum wheels and the sun and nadir sensors. In order to have a successful
mission that CPU must be able to reliably collect data about its attitude and be able to control its
pointing accuracy to < 1 degree. This pointing accuracy is necessary so that when burns are
completed SPDR sat can be sure that it will rendezvous with the target.
The second section that the CPU is in charge of is performing the burns that will take the
SPDR sat from its launch orbit to the desired orbit where it can collect space debris. To properly
make a burn the CPU must have an approximate locations of itself and the target debris. The
approximate location of the spacecraft will be determined via ground station gps information,
and the debris orbit information will be transferred by the same method. For this the CPU must
be able to communicate with the transceiver which will be getting information from the antenna.
With this information and control of the propulsion system SPDR sat will make a burn to reach
its target.
The final section for the CPU is the find and capture section. Here the SPDR sat must
locate the debris, approach it and then launch the capture device so that the mission can be
completed. One of the most difficult requirements is that the final rendezvous and collection
must be done autonomously. This means that SPDR will have to “see” the debris and know that
it wants to capture it with the capture device. In order to “see” the debris the team has decided
that a forward facing camera would be best. Once the debris has been recognized by the craft
SPDR sat will orient itself to face the debris, launch its capture mechanism and then launch a
drag sail that will decrease its deorbit time.
3.6 Communication Requirements
The communication system requires compatible antenna and transceiver components
capable of operating in the UHF, VHF or S-band frequency range as defined by the International
Telecommunication Union (ITU). The transmission power and receiver sensitivity of the
transceiver must be sufficient to establish a communication link with a given ground station from
LEO, and the uplink and downlink data rates must be sufficiently high (minimum 1200 bit/s) to
guarantee the necessary telemetry can be communicated throughout the duration of the mission.
3.7 Thermal Requirements
In order for SPDR Sat to operate correctly, it must be able to withstand the environmental
conditions of Low Earth Orbit. Using the International Space Station’s orbit altitude of 400 km
as a point of reference, temperatures range from -157 C to 121 C (NASA Science). These
temperatures are far beyond the operating temperatures, listed in Table 3.7.1, of the on-board
electrical components.

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Table 3.7.1: Minimum and maximum allowable temperatures for operation
Component Min. Temp (C) Max Temp (C)
Thruster 0 50
Cube Computer -25 60
Battery -10 40
Solar Panels -40 125
Antenna -30 70
Transceiver -20 60
Momentum Wheels -20 50
Sun Sensor -25 50
Debris Camera -20 60
For mission success the internal temperature of the CubeSat must be maintained at a
common temperature suitable for all components; consequently a thermal management system is
essential. Important is not only the temperature outside the spacecraft, but the length of the
spacecraft’s duration at the different environmental conditions. For this reason, the eclipse and
sunlit periods must be taken into account. Furthermore, a thermally optimal component
placement must be designed. Compromises in component placement and orientation are required
in order to satisfy CDS 3.2.14.1-4 which constrains the spacecraft’s center of mass. The thermal
control system will be designed using flight-tested, COTS hardware to ensure its parts adhere to
current launch and in-space regulations.
The SPDR Sat’s thermal management system will need to prioritize internal heat
retention over heat rejection. The spacecraft will likely undercool before it overheats based on
the thermal conditions in LEO. SPDR Sat’s thermal control solution must rely solely on
conduction and radiation as methods of heat transfer. The proposed design must consider thermal
effects of the Sun’s direct radiation, albedo, and planetary infrared radiation experienced at

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different locations along the orbit. Understanding the spacecraft’s orientation at those times is
also necessary for understanding which surfaces are exposed to solar radiation and which
experience radiation from Earth. Components requiring air-based convection should not be used
as they would not function in the vacuum environment present in space.
There are certain trade-offs which the thermal management system is faced with. When
contemplating whether to use passive or active thermal control, available power, mass
limitations, and budget constraints are substantial concerns. Essential, but heavy, power
consumers are the propulsion system, electronics, and reaction wheels. A battery can either be
chosen so that it can solely handle the power-dependent systems mentioned above while
maintaining a small margin for degradation, or it can be chosen to have enough excess power to
support an active thermal management system in addition to the other systems. Section 4.3 of
this report will discuss pros and cons of various options for thermal regulation.

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4 Subsystem Concept Development
4.1 Structural ConceptDevelopment
The main component of the structure is the frame of the CubeSat. The SPDR Sat uses the
RadiusSpace off-the-shelf frame because of its low mass (255 grams). Significant testing will
need to be conducted to ensure this frame will not break under the mission loads. The frame is
made of aluminum with titanium screws for strong connections. Other considerations include the
CubeSatShop Frame and the Princeton Satellite Systems Frame. Both of these were heavier than
the RadiusSpace frame. These will remain possibilities if the RadiusSpace frame cannot
withstand the loads, but it is expected that the RadiusSpace frame will be sufficient and pass all
testing.
Figure 4.1.1: RadiusSpace Frame (RadiusSpace)
When designing the deorbiting system, the team had two choices to make. How to attach
to the debris, and how to deorbit the debris once it is attached to the satellite. For the attachment
subsystem, the first few concepts that were investigated included: a mechanical grasper, a
magnetic system, and a net for catching debris. When considering the grasper, it was appealing
because it would be reliable and repeatable. Ultimately, it was decided that this option would not
be viable because a grasper would not be universal to all shapes of debris and thus would only be
useful for a very specific size and shape. In regards to a magnetic system, there was appeal in the
fact that it would not require any moving parts, and would not take up much space in an already
tightly space-constrained system. Potential problems with a magnetic system are that it would
not be able to attach to non-magnetic debris. Ultimately, the team decided on a net to catch the
debris because it would be able to catch any shape debris. Obviously, this system would not
work for catching smaller debris due to the holes in the net. The net would predominantly be
used for catching debris on the scale of 10 cm. or larger.
The CubeSat will not use commercial off the shelf hardware the capture device. The team
will create a capture device closely resembling the European Space Agency “fishing net” design,
but for a smaller satellite and a smaller piece of debris. Due to the European Space Agency’s
(ESA) success on this design, which has been tested in microgravity parabolic-arc test planes, a

21. 21
similar success can be expected for the design with only a reduction in scale. The concept for the
net design can be seen in Figure 1.1.1 of the ESA “fishing net”.
It will used compressed gas to launch weights attached to a net. The weights will wrap
around the debris and tangle themselves. The cord attached to the net will allow the CubeSat to
drag the debris back to earth using a drag sail.
In our design of the capture device, we will include the front facing camera system. The camera
is needed to face the debris during capture to ensure a successful rendezvous and capture. The
actual camera chosen, the ESTCube-1 camera, will be explained in the primary computing
section.
For the deorbiting system two main concepts were investigated. The first idea was to use
solely propulsion to lower the orbit of the desired space debris. While this is reliable and simple,
a problem that might be encountered using this method would be running out of fuel before the
mission was accomplished. In some cases, this would lead to a complete mission failure, and in
other cases would lead to the debris taking much longer times to deorbit. The method chosen was
a commercial off the shelf deployable drag sail. When the satellite is attached to the debris, the
drag sail will deploy to increase the surface area and lead to the debris deorbiting.
The team chose the Clyde Space Aerodynamic End Of Life Deorbit System for CubeSats,
or AEOLDOS. This drag sail was created specifically for debris prevention. It can be used for
any CubeSat mission for timely removal at end-of-life. This is the only current off-the-shelf drag
sail that is available specifically for end-of-life removal. The sail deployment takes no power
from the system. This will save the satellite on battery power and fuel needed for active deorbit.
Most importantly, the company lists that the sail can be attached mid body, which is greatly
desired for the SPDR Sat setup. Setup will be explained in the final design section.
4.2 PowerConceptDevelopment
4.2.1 – Battery Selection
When researching for COTS battery systems, there are a multitude of options from many
different suppliers. Each of these has different form factors, storage capacities, and prices.
Because of the high amount of options, some specific criteria had to be set so that the options to
choose between could be minimized. Listed in Table 4.2.1 below are different specifications of a
few of the batteries that were to be decided between. Below that is the selection matrix in Table
4.2.2 that was created between the different battery options.

22. 22
Table 4.2.1: Battery Specifications Sheet
Batteries Price
($)
Capacity
(Whr)
Weight
(grams)
Specific
Capacity
(Whr/g)
Cost
per
Whr
($/Whr)
Thickness
per
Capacity
(mm/Whr)
Thickness
(mm)
Thermal
Range
(degC)
Optimal
Ratios
Larger
=Better
Smaller
=Better
Smaller
=Better
Crystalspace
“Vasik”
5,900 11 80 0.1375 536 0.636 7 -10 – 40
2x “Vasik” 11,200 22 140 0.1571 509 0.591 13 -10 – 40
NanoPower
BP4
2700 38.4 240 0.16 70.31 0.599 23 Not
Listed
NanoPower
BPX
6000 77 500 0.154 77.92 0.527 40.6 -40 – 85
(heater)
CubeSat
BM1
Not
Listed
40 310 0.129 n/a 0.628 25.1 Not
Listed
CubeSat
Linear EPS
Not
Listed
22 210 0.1048 n/a 1.255 27.6 Not
Listed
Table 4.2.2: Battery Selection Matrix
“Vasik” 2x “Vasik” NanoPower
BP4
NanoPower
BPX
CubeSat
BM1
CubeSat
Linear EPS
Storage
Capacity
0 0 1 1 1 1
Thickness 1 1 0 -1 -1 -1
Mass 1 1 1 -1 -1 0
Price 0 0 1 0 n/a n/a
Integration
Ability
1 1 0 0 -1 -1
Specific
Capacity
0 1 1 1 0 0
Thickness
vs.
Capacity
0 1 1 1 0 -1
Aggregate 3 5 5 1 -2 -2
When one takes a look at the spec sheet in conjunction with the selection matrix, it can be seen
that the front running batter selections are the dual “Vasik” battery and the NanoPower BP4. In
choosing, though, special attention had to be drawn to the thickness and mass properties of each.
Both have similar thickness per capacity values, but the “Vasik” far outperforms in absolute
values. In fact, one could add an additional “Vasik” battery to the stack and still have a smaller
thickness and less mass than the BP4. Because of these two key features which are both
weighted very heavily in the design of a CubeSat, the dual “Vasik” battery was chosen to
provide storage. In section 11.2 simulations are run to verify that this battery choice has enough
capacity.

23. 23
4.2.2 – Solar Panel Selection
As stated in section 3.2, the main constraint focused on with solar panel choice was
whether or not they could recharge a depleted battery while on the sunlight side of an orbit.
Below are the solar panel spec sheet followed by the accompanying selection matrix.
Table 4.2.3: Solar Panel Specifications Sheet
Name Optimal
Power
Output (w)
Mass (g) Cell
Efficiency
Cost ($) Cost to
Output
Comparison
($/W)
Mass to
Output
Comparison
(g/W)
Optimal
Ratios
Smaller
=Better
Smaller
=Better
Clydespace
3U long-side
deployables
12.61 1,662 28.3% 92,700 7,351.3 131.80
Clydespace
3U short-side
deployables
15.03 1,662 28.3% 80,600 5,362 110.57
ISIS 1U
Panel
2.3 50 28% 2,734 1,189 21.74
Dual ISIS 2U
deployables
9.2 200 (not
including
hardware)
28% 10,934 1,188 21.74
NanoPower
Solar P110
2.3 59 28% 2180 947.8 25.65
Table 4.2.4: Solar Panel Selection Matrix
Clydespace
3U long-side
Clydespace
3U short-side
2x ISIS 2U
Deployable
ISIS 1U
Body
NanoPower
P110 1U
Body
Power Output 1 1 1 0 0
Mass -1 -1 1 1 1
Cost -1 -1 1 1 1
Cost to
Output
Comparison
-1 0 1 1 1
Mass to
Output
Comp.
0 0 1 1 1
Aggregate -2 -1 5 4 4
As can be seen, for the parameters chosen, the relative cost and mass of both of the Clydespace
solar setups far exceeds the others. The dual ISIS 2U deployable solar panel setup was chosen
because of the scalability of the individual cells along with its exact meeting of all criteria. The
output values are evaluated to be high enough in section 5.2.

24. 24
4.3 Propulsion ConceptDevelopment
There were many aspects that needed to be considered when choosing a propulsion
system. The first choice that had to be made was to determine whether or not the team wanted to
use a chemical propellant propulsion system, or one that utilized ion thrusting. Modern ion
thrusters mainly use inert gases for their propellant, which is injected from the downstream end
of the thruster and flows toward the upstream end. Electrons from the discharge cathode ionize
the propellant by means of electron bombardment. High-strength magnets are placed along the
discharge chamber walls so that as electrons approach the walls, they are redirected into the
discharge chamber by magnetic fields. By maximizing the length of time that electrons and
propellant atoms remain in the discharge chamber, the chance or ionization is maximized. The
thrust force is the force that exists between the upstream ions and the accelerator grid, which is a
downstream electrode that is charged highly negative. The exhaust velocity of the ions in the
beam is based on the voltage applied to optics (Dunbar). A very limited resource on a satellite, is
the power. The amount of electricity needed to deliver the needed thrust to change orbit
orientation in low earth orbit is the big reason why the team decided to stay away from ion
propulsion. Because of this, the team decided to look into chemical propulsion systems. The
three systems that were considered, along with two ion thrusters, can be seen in the table below:
Table 4.3.1: Propulsion Selection Matrix
Subsystem Delta-V Thrust Specific Impulse Mass
Power
Requirement
Total
Points
VACCO ADN MiPS 1 2 3 4 3 13
Aerojet MPS-120 2 1 4 5 2 14
Busek BIT-3 4 5 1 3 5 18
Nano Sat MiPS 3 4 5 2 1 15
Clyde Space Plasma Thruster 5 3 2 1 4 15
Scoring System: 1=Best 5=Worst
Between the three chemical propulsion systems, the given delta-v, thrust, specific impulse and
power required to operate the system, were are important engine parameters that the team graded
on a scale from one to five, one being the best and five being the worst. After this grading
system was complete, it was clear to the group that the VACCO ADN Micro Propulsion System
is the best propulsion system for this mission. With a high delta-v, specific impulse and thrust it
is the most well-rounded throughout those three categories. Although this system has a
relatively high mass and power requirement, it was determined that these values were ones that
would be able to fit within our mass and power budgets. The other benefit of the VACCO ADN
MiPS is the type of propellant it utilizes. The propellant is a green propellant based on
ammonium dinitramide (ADN). The most widely used liquid monopropellant used today is
hydrazine, well known for its good performance characteristics, but has limitations and liabilities
regarding toxicity, operational handling and environmental impact. ADN is a high-energy
inorganic salt, mainly intended as an oxidizer in solid rocket propellants. However, through

25. 25
developmental work done in the 1990’s, it was found that ADN was highly soluble in polar
solvents, which led to the realization that it could also be used as an oxidizer in liquid propellants
(Anders). A table displaying properties of both hydrazine and FLP-106 (liquid monopropellant
form of ADN):
Table 4.3.2: Properties of hydrazine and FLP-106
As can be seen from Table 4.3.2, the two propellants are very comparable in their properties,
with FLP-106 even having a slightly higher specific impulse, and being less toxic and volatile to
the environment. The only drawback to using an ADN based monopropellant is the high
combustion temperature. Although, there has been a lot of research going into this topic, with
three different methods being identified (Anders):
1. Pyrotechnic (by forming hot gases using a solid energetic material which in turn
will hear the propellant)
2. Thermal Conduction (by spraying the propellant on a hot object which in turn is
heated by electric means)
3. Resistive (ADN is a salt and the propellants thereby possess a relatively high
electric conductivity. This means that an ADN-based monopropellant can be
resistively heated)
To summarize, the VACCO ADN Micro Propulsion System was chosen for the SPDR’s
propulsion system for its ideal combination of delta-v, thrust, specific impulse and type of
propellant with the mass and power properties being able to fit within the mass and power
budgets.
4.4 : Attitude Determination and Control SystemConceptDevelopment
Research was done into which ADCS systems had the performance specifications that
went along with our mission requirements. Three COTS systems were investigated in this phase
of design.
Table below shows a comparison between the MAI-400 from Maryland Aerospace, the XACT
from Blue Canyon Technologies, and Cube ADCS from CubeSatShop. The product brochures
for each system are located in the appendix.

26. 26
Table 4.4.1: ACDS Performance Specifications
Product MAI-400 XACT Cube ADCS
Pointing Accuracy .2 degree .003-.007 degree .3 degree
Mass .6931 kg .85 kg .428 kg
Volume 10x10x5.59 cm 10x10x5 cm 10x10x5 cm
Electronics Voltage 5 V 5 V 5 V
Reaction Wheel
Voltage
12 V 12 V 5 V
Maximum Torque .625 mNm .6 mNm .23 mNm
Momentum Storage 11.8 mNms 15 mNms 1.7 mNms
Nominal Power Draw 1.5 W 1.87 W .5 W
Maximum Power
Draw
3.2 W 2.83 W 1.5 W
Data Interface RS-232 RS-442, I2C, SPI I2C, UART
Operating
Temperature
-20C to 60C -30C to 70C -10C to 70C
Additional Comments Integrated flight
computer for
propulsion and
communication
control
Cost $36,995 $125,000 $28,750
Since not all specifications for the ADCS are equally important, a weighted selection
matrix was used to determine which product to pursue. Table 4 shows this selection matrix.
Table 4.4.2 ADCS Selection Matrix
Product Weighting MAI-400 XACT Cube ADCS
Pointing Accuracy 2 0 1 0
Mass 3 0 0 1
Volume 3 1 1 1
Electronics Voltage 2 1 1 1
Reaction Wheel
Voltage
2 0 0 1
Maximum Torque 2 1 1 0
Momentum Storage 2 1 1 0
Maximum Power
Draw
2 0 0 1
Data Interface 1 0 0 1
Operating
Temperature
1 1 1 1
Additional
Comments
2 0 0 1
Cost 3 1 0 1
Weighted Total 13 12 19

27. 27
When making the selection matrix, the most important factors included cost, volume and
mass. Other important factors which helped to make the decision were data interface and
additional comments. The section on additional comments included the fact that the Cube ADCS
can incorporate a full satellite computer system to control propulsion and communications
through the ADCS computing stack. This factor ultimately made the difference in choosing the
Cube ADCS because it was cheaper and helped the team save on space and mass because all of
the satellite computing could be done aboard the ADCS system. While the other two systems had
superior torque and momentum capabilities, their excessive power draw requirements and poor
computing interface specifications made them bad choices for the SPDR Sat design. A Computer
Aided Design (CAD) sketch of the system is shown in Figure below.
Figure 4.4.1: Cube ADCS(CubeSatshop)
From the brochures associated with the Cube ADCS, many control modes are available such as
B-dot, Y-spin, and Wheel momentum Cross product Control. Additionally, included in the
ADCS library that comes with the system are a variety of Kalman Filters for attitude
determination, including a Robust Rate one for initial de-tumbling of the CubeSat.
4.5 Primary Computing Concept Development
Found on the next page in table 4.5.1 are the three CPU options that were considered for
this mission.

28. 28
Table 4.5.1: CPU Considerations
Name Cube Computer Q6 Clyde-Space Mission
Computer
Power <200mW 1W 1.25W
Mass 70g 23g 94g
Size 96 x 90 x 10mm 78 x 38 x 19mm 96 x 90 x 12mm
Op Temps -10C to +70C -40C to +85C -25C to +65C
Interface I2C, UART, CAN RJ45, Rs-232, I2C, SPI, Serial UART
Price $4960.00 $20675.00 $5240.00
Ram 64MB 256MB 64MB
Mem Storage up to 16GB up to 8GB 2GB
For the selection matrix, each of the CPU options were ranked on a variety of characteristics that
were considered important for the mission (the matrix ranks the CPU 1 for being the best in that
category and 3 for being the worst). This selection matrix can be seen in table 4.5.2 below.
Table 4.5.2: CPU Selection Matrix
Selection matrix Cube Computer Q6 Clyde-Space Mission
Computer
Thermal Range 3 1 2
Size 2 1 3
Memory 2 1 3
Price 1 3 2
Mass 2 1 3
Total(lower =
better)
10 7 13
From the selection matrix it can be seen that the two best options are either the Q6 or the
CubeCompter. At this point in the selection process the focus falls on cost and interface type.
After reviewing a variety of different components and their interface connections it was found
that a CPU with and I2C or a UART connection would be best because many components have

29. 29
this interface. With this in mind the CubeComputer was chosen because it was much cheaper
than the Q6, it also has both the I2C & UART interfaces.
4.6 Communications SystemConceptDevelopment
There was not a wealth of available hardware for the communications subsystem. The
utilization of an omnidirectional antenna was determined a priori, so selection of COTS
transceivers was limited to those operating in UHF/VHF frequencies, as the available S-band
antennas are not omnidirectional. Therefore, selection of a transceiver came down to three main
choices. The first was the NanoCom U482C UHF Half-duplex Transceiver, displayed in Figure
4.6.1.
Figure 4.6.1: NanoCom U482C UHF Half-duplex Transceiver(CubeSatshop)
This transceiver operates in UHF frequencies (435-438MHz), with an uplink data rate of 1200-
4800 baud, downlink data rate of 1200-9600 baud and transmit power of ~27dBm (~500mW). It
is half-duplex meaning that communication is only possible in one direction at a time (i.e. you
cannot be transmitting and receiving simultaneously). Its peak power requirement during
transmission is 5500mW, has dimensions 95.40 x 90.15 x 18.00 (mm) and weighs 75g, all for a
price of ~$8500. (CubeSatshop)
The second option was the ISIS VHF downlink / UHF uplink Full Duplex Transceiver,
shown in Figure 4.6.2.

30. 30
Figure 4.6.2: ISIS VHF downlink / UHF uplink Full Duplex Transceiver(CubeSatshop)
This transceiver operates in both UHF and VHF frequencies, transmitting in the 130 – 160 MHz
range and receiving in the 400 – 450 MHz range. It has an uplink data rate of 300-1200 baud, a
downlink data rate of 1200-9600 baud and a transmit power of ~22dBm (159mW). It is full
duplex meaning that communication in both directions is possible simultaneously. Its peak
power requirement during transmission is 1.7W, has dimensions 96 x 90 x 15 (mm) and weighs
85g, sitting at a cost of ~$9000. (CubeSatshop)
The final consideration was the ISIS UHF downlink / VHF uplink Full Duplex
Transceiver.
Figure 4.6.3: ISIS UHF downlink / VHF uplink Full Duplex Transceiver(CubeSatshop)
Similar to the last product, this transceiver also operates in both UHF and VHF frequencies;
however, its uplink and downlink frequencies are switched, meaning this transceiver transmits in
the 420-450MHz range and receives in the 140-150MHz range. It has an uplink data rate of
1200 baud, a downlink data rate of 1200-9600 baud and a transmit power of ~27dBm (500mW).
Similar to the last one, it offers the advantage of being a full duplex transceiver, allowing for
simultaneous transmission and reception of communication signals. Its peak power requirement

31. 31
during transmission if 4.0W, has dimensions 96 x 90 x 15 (mm) and weighs 75g, sitting at a cost
of ~$9000. (CubeSatshop)
In order to select the best option from the above three components, the most important
parameters given above were compared in a concept selection matrix, shown below.
Table 4.6.1: Concept Selection Matrix for Transceiver
Option 1 Option 2 Option 3
Light-Weight + - +
Power-efficient - + -
Small size - 0 0
Powerful transmission + - +
Simultaneous Comm. - + +
Sufficient Data
Transfer
+ - 0
Total 0 -1 2
Since Option 3 scored the highest, it was selected for the SPDR Sat. It should also be noted that
it only received a – in the power-efficient criteria, and that with the current design the SPDR Sat
has a large power margin of operation.
The next piece of hardware to select was the antenna. With the transceiver already
selected, the goal was to target antennas designed for transmission in the UHF frequency that can
also receive VHF signals. A second required criteria was that the antenna be omnidirectional –
so as to save propellant by not treating orientation for the sake of communication as a primary
concern. These two criteria made the selection process simpler, with only one antenna system
really standing out. Other antenna considerations operated exclusively in the S-band (such as the
CPUT S-band Patch Antenna from Clyde(Clyde-space) ) or in UHF frequencies (such as the
ANT430 UHF Turnstile Antenna from NanoCom(Clyde-space)). The selected antenna is the
ISIS Deployable Antenna System for CubeSats(Figure 4.6.4), which is compatible with any UHF
and/or VHF radio system.
Figure 4.6.4: ISIS Deployable Antenna System(CubeSatshop)

32. 32
The system can be configured in 5 different ways: as 4 VHF or UHF monopoles, as a single
VHF or UHF dipole, as dual VHF or UHF dipoles, as VHF or UHF turnstile or as a combination
of dipole and monopoles. Its maximum power consumption is 2W during deployment, with a
nominal power consumption of 20mW. It has a useable bandwidth of 10MHz in the relevant
frequency range (be it UHF or VHF), with a return loss of -10dB and beam gain is given as 0dBi.
The system has a mass of <100g, size of 98 x 98 x 7 (mm) and costs ~$4775. Furthermore, there
is a 30mm diameter center-hole in the system that could fit a camera or some other payload if
convenient. (CubeSatshop)
4.7 Thermal ConceptDevelopment
Development of the thermal management system begins with understanding the methods
of heat transfer which the spacecraft will experience.
Convection may be ignored when performing a thermal analysis of SPDR Sat. This is
because SPDR Sat will be operating in space which is considered a vacuum.
Conduction will play a large role in temperature regulation of the spacecraft. It will be
one of the tools used to remove excess heat generated by electronics. The diagram and equation
in Figure 4.7.1 below show this relationship.
Figure 4.7.1: Conduction between two surfaces(Brighthub)
In Figure 4.7.1, q represents rate of conduction heat transfer, k is thermal conductivity, A
is cross sectional area, T1 and T2 represent the temperatures on each surface, and L is distance
between the two points of reference. In thermal analysis these reference points are known as
“nodes” and they define the capacitance of a particular region. Each structure analyzed is divided
into a finite number of regions known as “subvolumes” to simplify the analysis. In addition to
the effect of temperature differences on conduction, the temperature magnitudes also have an
effect. The higher the temperatures experienced, the greater the material’s conductivity and thus
the greater its conductance. The converse also holds true. The driving force for conduction in the
CubeSat will be those between the higher temperature electronics transferring heat to the cooler
aluminum frame and to the more thermally resistive circuit boards. In order to increase the rate at
which heat is removed from the electronics to prevent overheating, a solution relying on
increased surface area or increased conductivity would be effective. An example of this would be

33. 33
a liquid cold plate heat sink, depicted in Figure 4.7.2, which is a metallic device possessing a
high thermal conductivity and high surface area to volume ratio.
Figure 4.7.2: Liquid cold plate heat sink
A less ideal solution would be to increase wall thickness as this would require more space
and add to the limited mass budget. Preventing undercooling via conduction could be achieved
either through a resistive heater on a circuit board or thermal insulation blanket. These solutions
would attempt to maintain the temperature difference between the hot and cold regions at an
acceptable level. They would effectively produce a heat lost rate small enough to prevent the
temperature of the electronics from reaching critically low levels throughout the eclipse period.
Useful metrics for quantifying the performance of various heat transfer solutions are
thermal conductivity and resistance to conductive heat transfer, Rcond. Resistance to conductive
heat transfer is inversely proportional to the thermal conductivity and for thermal analyses it can
be treated similar to a resistor in an electrical circuit. Equations 4.7-1 & 4.7-2 below are used in
providing a detailed quantitative analysis of the proposed designs.
𝑆𝑒𝑟𝑖𝑎𝑙: 𝑅𝑡𝑜𝑡𝑎𝑙 = 𝑅1 + 𝑅2 + ⋯ + 𝑅 𝑛 (4.7 − 1)
𝑃𝑎𝑟𝑎𝑙𝑙𝑒𝑙: 𝑅𝑡𝑜𝑡𝑎𝑙 =
1
𝑅1
+
1
𝑅2
+ ⋯ +
1
𝑅 𝑛
(4.7 − 2)
An effective insulator used today is FR-4; this material is used in circuit boards. This
composite material has a high resistance to thermal conduction which is one of the reasons why
it is widely used with electronics.
Radiation is another important phenomenon to consider for the thermal analysis of SPDR
Sat. The following equation shows radiation is proportional to the difference of temperatures to
the 4th
power.
𝑃 = 𝑒𝜎𝐴( 𝑇4
− 𝑇𝐶
4 ) (4.7 − 3)
In Equation 4.7-3, P represents the net radiated power, e is the emissivity, 𝜎 Stefan’s
constant, A is radiating area, T is temperature of radiator, and TC is the temperature of the

34. 34
surroundings (AAU Student Space). Radiation from component surfaces can most easily be
adjusted through manipulating the emissivity of the surfaces. Ideally the spacecraft would have a
variable emissivity where it could be a black body during eclipse and white body during sunlight.
This would help it absorb and retain heat at its coldest point in orbit and reject heat during the
hottest orbit location. Unfortunately, true black and white bodies are difficult to obtain, thus
constant grey bodies are a more realistic design expectation. Since cold temperatures will be a
greater concern than overheating, the emissivity of surfaces should be designed with a cold bias
where more emphasis is placed on absorbing heat. This can be accomplished using thermal
coatings sprays and/or multilayer insulation. Figure 4.7.3 shows the materials making up the
various layers in multilayer insulations.
Figure 4.7.3: Structure of MLI blanket including common materials in each layer
As seen in Figure 4.7.3, the multilayer insulation is composed of about 3 core materials, forming
repeating layers. Usually, between 2 and 35 insulating layers produce a sufficiently effective
radiation barrier. The thickness of each of these layers varies by material, but all are roughly on
the order of 6 micrometers.
In an effort to streamline the task of selecting components for the thermal system, a
selection matrix was created. The selection matrix below uses key metrics to compare the
performance of the heat transfer devices under consideration.

35. 35
Table 4.7.1: Thermal Component Selection Matrix
Passive Active
Multilayer
Insulation
(MLI)
Thermal
Coating
Spray
(Inorganic
Optical
Black)
Radio
Isotope
Heater
Liquid Cold Plate
Heat Sink
Polymide
Thermofoi
l Heater
Mass 94.60g Negligible 39g Unknown Unknown
Cost Unknown Unknown Unkno
wn
Unknown Unknown
Size 20 cm x
40 cm
(thickness:
0.24 cm)
(Thickness:
0.0635 mm)
2.54
cm x
3.30cm
5.72 cm x 13.34 cm x
1.52 cm
12.7 mm x
50.8 mm
Power
Required
0 W 0 W 0 W > 0 W 2.52 W
Temp
Operating
Range
-180 to
+150 C
-180 to 600
C
Unkno
wn
Unknown -200 to
+200 C
Emissivit
y
.75 0.91+/-0.02 Unkno
wn
Unknown Unknown
From the selection matrix it can be seen that the devices are broken up into two categories, active
and passive. Passive systems have 0 Watt power requirements while the active systems require a
nonzero power and controller to operate. Since power is limited, requiring power to operate is
often a drawback. Power distribution primarily allocated to running the CPU, ACS, and
propulsion system. The remaining power may not be sufficient to power an active thermal
device, and introducing a higher capacity battery would increase the mass of the spacecraft and
possibly not fit. The negative consequences of selecting a higher capacity battery outweigh the
benefits. Also, the active thermal control systems are larger in size than the passive ones. This is
a major negative for active systems, again, due to the limited space and limited allowable mass.
Although cost of these components is not currently known, it can already be seen that the
negative aspects associated with active systems, make passive systems the superior design
choice.

36. 36
5 Design Analysis
5.1 Structures SystemAnalysis
The table below shows the SPDR Sat parts with their associated sizes and masses.
Table 5.1.1: Part List for Mass and Volume Analysis
System Part Mass (kg) Size (cm)
Capture Device Proprietary Net Launcher ~.4 ~10 x 10 x 5.5
Frame 3U (Radius Space) 0.255 10 x 10 x 30
Propulsion
Vacco ADN Micro Propulsion
System (VACCO)
1.8
10 x 10 x 10.55
(Plus Thrusters)
CPU
On-Board Momentum Wheel -
Cube Computer (Cube
Computer)
Included in
Cube ADCS
Included in ADCS
(9.6 x 9 x 1)
Battery
Dual Crystalspace P1U
“Vasik” (Crystalspace) [x2]
0.14 9.6 x 9 x 1.8 [x2]
Solar Panels 2U (ISIS CubeSat) [x2] 0.200 Deployable (0.2)
Drag Sail AEOLDOS (AEOLDOS) 0.372 10 x 10 x 4
Communications
1: Deployable Antenna System
(Deployable)
2: Full Duplex Transceiver
(ISIS UHF)
1: <0.1
2: 0.075
9.6x9x1.5
Momentum Wheel
(includes CPU specs)
Cube ADCS (Cube ADCS) 0.316 <10x<10x4.8
Sensor ESTCube-1 Camera 0.03
Included in
Capture System
Thermal
1: Multi-layer Insulation
2: Thermal Coating
1. ~0.145
2.~0
1: ~0
2: ~0
TOTALS 3.833 kg
29.96 cm
The SPDR Sat is expected to have a total mass of 3.833 kilograms. This allows for a buffer of
4.4%. This buffer may reduce as the parts can come in slightly overweight. Also, some of the
parts only have approximate masses. The mass buffer should allow the satellite to remain under
its restraining value of 4 kilograms even if all parts come in slightly overweight.

37. 37
The SPDR Sat stack height comes out as 29.96 centimeters. This is just below the required value
of 30 centimeters, however, this allows for near maximum size of the capture device design. If
there are changes to the design height, the capture device can be scaled down to account for the
change.
In the original design, the satellite allowed 3 centimeters in stack height for the capture
device. This was expected to be suitable for catching a piece of debris of size 30 centimeters in
all directions. With the now allowed 5.5 centimeters, the capture device is expected to be capable
of capturing a piece of debris 50 centimeters in all directions.
The drag sail opens to an area of 1.5 square meters. The initial area of the debris expected to
collect is .25 square meters maximum. Furthermore, the drag of the debris should be much less,
depending on the shape and drag coefficient, than a flat plate of 0.25 square meters. The drag on
a body is proportional to the coefficient of drag and the effective area of the face.
𝐷 = 𝐶 𝑑
1
2
𝜌𝑣2
𝐴 (5.1 − 1)
If we assume the debris is a cube at maximum size, with a drag coefficient 0.8 to1.05 (depending
on angle) and a face area of 0.25 square meters, drag will be proportional to approximately 0.116
multiplied by density and velocity squared. With the drag sail deployed, the drag increases to
1.92 multiplied by density and velocity squared. (Cd flat plate = 1.28 and A=1.5m2
) The drag
force increases to 8.3 times its original value. The deorbit time will decrease significantly due to
this increased drag.
5.2 PowerSystems Analysis
In the design of the system, an overall amount of power needed to power all subsystems
had to be established. In the initial design, it was desired to be able to run all components
simultaneously for an indefinite amount of time. Later in analysis, though, it was found that
although this capability was needed, the indefinite part was not. As a result of this, typical power
setups that may be used through the different stages of the mission were developed, as listed on
the next page in Table 5.2.1.

38. 38
Table 5.2.1: Various Power Setups
Phase/
Part Name
P-Pod
Launch
Pre-Nominal
Orientation
Nominal
Operation
Mission
Execution
End of Life
Cube Computer Yes Yes Yes Yes Intermittent
Cube Control Yes Yes Intermittent Yes Intermittent
Cube Wheels
(x3)
Yes Yes Intermittent Yes Intermittent
Cube Torquer
(x2)
Yes Yes Yes Yes Intermittent
Cube Sense No Yes Intermittent Intermittent Intermittent
Vacco ADN
Micro Prop.
No No No Yes No
Camera No No No Yes No
Antenna No Intermittent Intermittent Yes Intermittent
Transceiver No Intermittent Intermittent Yes Intermittent
Solar Panels Yes Yes Yes Intermittent Yes
Battery Yes Yes Yes Yes Yes
In the above table, a ‘yes’ indicates that the component will be running almost constantly,
an ‘intermittent’ indicates occasional use, and a ‘no’ indicates a lack of use through the execution
of the phase. From a quick look at the table also shows the variety of power setups that could
have been accounted for and designed to. But the easiest and most effective during validation of
product choice is to design towards both max power draw as well as the nominal power draw.
For the max power draw, it was chosen to be able to execute this through at least one full orbit.
The nominal power case, though, needed to execute the aforementioned indefinite execution.
This Indefinite execution indicates that the battery would supply enough power through the
eclipse and the solar panels would be able to recharge this partially depleted battery during the
sunlight portion of the orbit. Listed in Table 5.2.2 are these two cases and their required draw.
Table 5.2.2: Total Power Consumption Average
Max Power Setup Nominal Setup
Cube Computer 310 mW 170 mW
Cube Control 150 mW 150 mW
Cube Wheels (x3) 3.3 W 2 W
Cube Torquer (x2) 409 mW 409 mW
Cube Sense 360 mW 90 mW
Vacco ADN Micro
Propulsion
15 W n/a
Camera 280 mW n/a
Antenna 2 W 20 mW
Transceiver 4 W 480 mW
Total 25.809 W 4.309 W
As can be seen, the two different setups have very different power consumption needs.
Many of the components that were chosen were chosen for their low power consumption during

39. 39
nominal deployment and operation. Further analysis would need to be done to find out what
percentage of the orbit would be in both sunlight, penumbra, and umbra. Simulations were run in
STK to find these such things and whether these times would change through the mission cycle.
Two time periods were analyzed for a more accurate analysis pertinent to the proposed mission.
The first time period chosen was 04 July 2017 to 05 July 2017. The second chosen was 5 months
later from 04 December 2017 to 05 December 2017. Two separate orbits were also analyzed for
each of these time periods. Both orbits were circular and at 400 and 600 km in altitude.
Show below is a visual representation of one of such analyses completed in finding the
lighting points. For this graph, the time length was shortened so as to allow for a better
representation of the data. Listed in Appendix E are the data table procured for each lighting
simulation executed.
.
Figure 5.2.1: Lighting Times at 400 km Orbit
Looking at Figure 5.2.1 it was determined that the orbit, as expected, has periodic light
(and dark) times. The longest time spent in full eclipse was experienced during the simulation
was 2138 seconds with an additional 18 seconds spent in penumbra. The accompanying time
spent in sunlight was 3386 seconds. The date this was experienced at was on 04 December 2017.
Because this date would put the most strain on the power supply, this was the date focused on
that the system would need to satisfy.
To find the amount of energy consumed during eclipse, the total amount of energy
needed (in watts) was taken and multiplied by the time spent in eclipse. This would convert the
number to joules which could be directly compared to the storage of the battery. The amount of
energy in the battery was found by multiplying its capacity (in watt-hours) by 3600 to convert to
joules. If there was still charge left in the battery, the max draw constraint would be satisfied.
 Total draw on system – 25.809 watts
 Time spent in eclipse – 2138 seconds
o Total draw – 55179 Joules
 Total storage of battery – 22 W-hr
o Total storage – 79200 Joules
 Excess storage – 24021 Joules
o Factor of safety – 1.4
For the nominal constraint, the same draw on the battery was used. The amount used by
the components would then be subtracted from the total storage of the battery. The amount
subtracted would need to be replenished completely while in the sunlight portion of the orbit.
There is more storage than draw. Therefore
all components can be run simultaneously
through eclipse while powered.

40. 40
The wattage provided by the solar panels would need to be multiplied by the time spent in
sunlight to find the number of joules that could be provided to the system. The optimal condition
is used because during orbit on the sunlight side, ADCS will be used to keep the solar panels at
the optimal 90 degree angle.
 Nominal draw on system – 4.309 watts
 Time spent in sunlight – 3386 seconds
 Time spent in eclipse – 2138 seconds
o Total draw – 23802.9 Joules
 Power output of solar panels – 9.2 watts
o Output in sunlight – 31151 Joules
 Output versus input – Excess of 7348
Joules
o Factor of safety – 1.31
As shown through the above calculations, the power system performs admirably despite
its somewhat lower performance when compared to more expensive products. But, combined the
two components selected will be able to provide electricity to all components necessary during
either the max power case or the more likely nominal case. Additionally, both situations have a
factor of safety of at least 1.3. Although it may seem that there is not enough power storage
because the time analyzed is only during one half of the orbit (eclipse) during the max power
case, there is not time during the mission during which the highest drawing systems (propulsion
namely) will be used constantly through an orbit. Because of this reality, the above listed
components satisfy the necessary constraints while being at both a low price and small form
factor and are thus the most desirable components to choose from.
5.3 Propulsion Systems Analysis
Analysis into the type of maneuvers the SPDR satellite may encounter was completed in
MATLAB. This was done to determine the amount of delta-v and propellant consumed for
various maneuvers, such as a Hohmann transfer, a phasing maneuver or an inclination change.
All MATLAB Code can be found in Appendix A.
5.3.1 Hohmann Transfer (Brown)
i. The velocity of the initial and final orbit, 𝑉, is expressed as
𝑉 = √
𝜇
𝑟
(5.3-1)
where 𝜇 is the planetary constant for earth, and 𝑟 is the radius of the initial and final orbits.
ii. The semi-major axis of the transfer ellipse, 𝑎, is expressed as
𝑎 =
𝑟𝑖 + 𝑟 𝑓
2
(5.3-2)
where 𝑟𝑖 is the radius of the initial orbit and 𝑟𝑓 is the radius of the final orbit.
iii. The velocity at periapsis of the transfer ellipse, 𝑉𝑝𝑡 , is expressed as
𝑉𝑝𝑡 = √
2𝜇
𝑟𝑖
−
𝜇
𝑎
(5.3-3)
The output of the solar panels during
sunlight is greater than the draw of the
nominal setup during combined eclipse
and sunlight times.

41. 41
iv. The velocity at apoapsis, 𝑉𝑎𝑡 , is expressed as
𝑉𝑎𝑡 = √
2𝜇
𝑟 𝑓
−
𝜇
𝑎
(5.3-4)
v. The velocity change required to enter the transfer orbit, Δ𝑉1, is expressed as
Δ𝑉1 = 𝑉𝑝𝑡 − 𝑉𝑖 (5.3-5)
vi. The velocity change required to “circularize” at final orbit, Δ𝑉2, is expressed as
Δ𝑉2 = 𝑉𝑎𝑡 − 𝑉𝑓 (5.3-6)
vii. The total velocity change, Δ𝑉𝑇, is expressed as
Δ𝑉𝑇 = Δ𝑉1 + Δ𝑉2 (5.3-7)
viii. The amount of propellant needed, 𝑚 𝑝, is expressed as
𝑚 𝑝 = 𝑚 𝑖 [1 − 𝑒𝑥𝑝(−
Δ𝑉 𝑇
𝑔 𝑐 𝐼𝑠𝑝
)] (5.3-8)
where, 𝑚 𝑖, is the initial mass of the satellite, 𝑔𝑐 is the gravitational constant and 𝐼𝑠𝑝 is the
specific impulse of the propellant (ADN).
For this maneuver, the following initial conditions and propulsion parameters were determined:
 𝑟𝑖 = 160 𝑘𝑚
 𝜇 = 398600.4
𝑘𝑚3
𝑠2
 𝑚 𝑖 = 4 𝑘𝑔
 𝑔𝑐 = 0.00980665
𝑘𝑚
𝑠2
 𝐼𝑠𝑝 = 258.81 𝑠
Table 5.3.1: Hohmann Transfer Propulsion Analysis
Propulsion Analysis for Hohmann Transfer
Altitude Change (km) Propellant Mass needed (g) Delta-V Required (km/s)
to 200 37.2966 0.0238
to 300 127.5904 0.0823
to 400 213.8787 0.1395
to 500 296.4014 0.1954
to 600 375.3809 0.2501
to 700 451.0228 0.3036

42. 42
Figure 5.3.1: Propellant Consumption vs Radius for Hohmann Transfer
Figure 5.3.2: Delta-V Requirement vs Radius for Hohmann Transfer
5.3.2 Phasing Maneuver (Curtis)
i. The initial orbit’s angular momentum, ℎ1, is expressed as
ℎ1 = √2𝜇√
𝑟 𝐴 𝑟 𝐶
𝑟 𝐴+ 𝑟 𝐶
(5.3-9)
where 𝜇 is the planetary constant for earth, 𝑟𝐴 is the perigee radius and 𝑟𝐶 is the apogee radius

43. 43
ii. The angular momentum for orbit 2, ℎ2, is expressed as
ℎ2 = √2𝜇√
𝑟 𝐴 𝑟 𝐷
𝑟 𝐴 + 𝑟 𝐷
(5.3-10)
where 𝑟𝐷 is the apogee radius of orbit 2 and is expressed as
𝑟𝐷 = 2𝑎2 − 𝑟𝐴 (5.3-11)
where 𝑎2 is the semi major axis of orbit 2 and is expressed as
𝑎2 = (
√ 𝜇𝑇2
2𝜋
)
2
3⁄
(5.3-12)
where 𝑇2 is the period of phasing orbit 2 and is expressed as
𝑇2 = 𝑇1 −
𝑡 𝐴𝐵
𝑛
(5.3-13)
where n is the revolutions of phasing orbit, 𝑇1is the period of orbit 1 and 𝑡𝐴𝐵 is the flight time
from the perigee A of orbit 1 to point B and are expressed as
𝑇1 =
2𝜋
√ 𝜇
𝑎1
3 2⁄
(5.3-14)
𝑡𝐴𝐵 =
𝑇1
2𝜋
( 𝐸 𝐵 − 𝑒1 𝑠𝑖𝑛𝐸 𝐵) (5.3-15)
where 𝑎1 is the semimajor axis of orbit 1, 𝐸 𝐵 is the eccentric anomaly of B and 𝑒1 is the
eccentricity of orbit 1 and are expressed as
𝑎1 =
1
2
( 𝑟𝐴 + 𝑟𝐶 ) (5.3-16)
𝐸 𝐵 = 2𝑡𝑎𝑛−1
(√
1− 𝑒1
1+ 𝑒1
tan
𝜃 𝐵
2
) (5.3-17)
𝑒1 =
𝑟 𝐶− 𝑟 𝐴
𝑟 𝐶+ 𝑟 𝐴
(5.3-18)
where 𝜃 𝐵 is the true anomaly of B.
iii. The speed at the perigee of orbit 1, 𝑣𝐴1 , is expressed as
𝑣𝐴1 =
ℎ1
𝑟 𝐴
(5.3-19)
iv. The speed at the perigee of orbit 2, 𝑣𝐴2 , is expressed as
𝑣𝐴2 =
ℎ2
𝑟 𝐴
(5.3-20)
v. The velocity change required to drop into the phasing orbit 2, Δ𝑉𝐴1, is expressed as
Δ𝑉𝐴1 = 𝑣𝐴2 − 𝑣𝐴1 (5.3-21)
vi. The velocity change required to return to orbit 1, Δ𝑉𝐴2, is expressed as
Δ𝑉𝐴2 = 𝑣𝐴1 − 𝑣𝐴2 (5.3-22)
vii. The total delta-v to complete the phasing maneuver, Δ𝑉𝑡𝑜𝑡𝑎𝑙, is expressed as
Δ𝑉𝑡𝑜𝑡𝑎𝑙 = |Δ𝑉𝐴1| + |Δ𝑉𝐴2 | (5.3-23)

44. 44
For this maneuver, the following initial conditions and propulsion parameters were determined:
 𝑟𝐴 = 160 𝑘𝑚
 𝜃 𝐵 = 45°
 𝜇 = 398600.4
𝑘𝑚3
𝑠2
 𝑚 𝑖 = 4 𝑘𝑔
 𝑔𝑐 = 0.00980665
𝑘𝑚
𝑠2
 𝐼𝑠𝑝 = 258.81 𝑠
Equation (5.3-8) was used to determine the mass of propellant required for this maneuver.
Table 5.3.2: Propulsion Analysis for Phasing Maneuver
Propulsion Analysis for Phasing Maneuver
Apogee Radius
(km)
Number of Rotations in Phasing
Orbit
Propellant Mass Needed (g) Delta-V Required (km/s)
200
2 506.3366 0.3435
3 337.9259 0.2240
4 253.5568 0.1662
5 202.8931 0.1321
6 169.1019 0.1096
300
2 494.1496 0.3346
3 329.7185 0.2183
4 247.3722 0.1620
5 197.9319 0.1288
6 164.9602 0.1068
400
2 482.3713 0.3261
3 321.7910 0.2128
4 241.4001 0.1579
5 193.1420 0.1256
6 160.9619 0.1042
500
2 470.9843 0.3179
3 314.1311 0.2075
4 235.6311 0.1540
5 188.515 0.1225
6 157.1005 0.1016

45. 45
Figure 5.3.3: Propellant Consumed vs Revolutions in phasing orbit
Figure 5.3.4: Delta-V requirement vs Revolutions in phasing orbit

46. 46
5.3.3 Inclination Change (Braeunig)
i. The change in velocity to change satellites orbital plane, Δ𝑉𝑜, is expressed as
Δ𝑉𝑜 = 2𝑉𝑖 𝑠𝑖𝑛
𝜃
2
(5.3-24)
where 𝜃 is the desired angle change and 𝑉𝑖 is the initial velocity, which is expressed as
𝑉𝑖 = √
𝜇
𝑟𝑖
(5.3-25)
where 𝜇 is the planetary constant for earth and 𝑟𝑖 is the initial radius of the orbit.
For this maneuver, the following initial conditions and propulsion parameters were determined:
 𝜇 = 398600.4
𝑘𝑚3
𝑠2
 𝑚 𝑖 = 4 𝑘𝑔
 𝑔𝑐 = 0.00980665
𝑘𝑚
𝑠2
 𝐼𝑠𝑝 = 258.81 𝑠
Equation (5.3-8) was used to determine the mass of propellant required for this maneuver.

47. 47
Table 5.3.3: Propulsion Analysis for Inclination Change
Propulsion Analysis for Inclination Change
Altitude
(km)
Inclination Change
(deg)
Propellant Mass
(g)
Delta-V
Required(km/s)
160
1 209.1074 0.1362
2 407.2686 0.2725
3 595.0435 0.4087
4 772.9648 0.5449
5 941.5390 0.6811
200
1 208.4875 0.1358
2 406.0937 0.2717
3 593.3732 0.4075
4 770.8540 0.5433
5 939.0385 0.6790
300
1 206.9619 0.1348
2 403.2009 0.2696
3 589.2597 0.4044
4 765.6545 0.5392
5 932.8770 0.6739
400
1 205.4695 0.1338
2 400.3702 0.2676
3 585.2327 0.4014
4 760.5623 0.5352
5 926.8404 0.6690
500
1 204.0093 0.1328
2 397.5993 0.2657
3 581.2893 0.3985
4 755.5739 0.5313
5 920.9245 0.6641
600
1 202.5800 0.1319
2 394.8861 0.2638
3 577.4265 0.3956
4 750.6857 0.5275
5 915.1252 0.6593

48. 48
Figure 5.3.5: Propulsion used vs. Inclination change
Figure 5.3.6: Delta-V requirement vs Inclination change
5.3.4 Desaturation of Momentum Wheels (Brown)
i. The propellant required to unload one momentum wheel, 𝑚 𝑝, is defined as
𝑚 𝑝 =
𝐻
𝐿𝐼𝑠𝑝
(5.3-26)

49. 49
where 𝐻 is the momentum of the wheel, 𝐿 is the momentum arm and 𝐼𝑠𝑝 is the specific impulse
of the propulsion system.
ii. The time required to unload the momentum wheel, 𝑡, is defined as
𝑡 =
𝐻
𝑛𝐹𝐿
(5.3-27)
where 𝑛 is the number of thrusters being utilized and 𝐹 is the force of one thruster
For this operation, the following initial conditions and propulsion parameters were determined:
Table 5.3.4: Momentum Arm Thruster Force
Direction Momentum Arm (m) Thruster Force (N)
Y 0.045 0.1
Z 0.045 0.0349
 𝐻 = 0.0017 𝑁𝑚𝑠
 𝐼𝑠𝑝 = 259 𝑠
 𝑛 = 2
MATLAB code that can be found in Appendix A: Propulsion Code was run and the following
values were obtained:
Figure 5.3.7: Matlab Output
5.4 ADCS Systems Analysis
5.4.1 Stability and Control
The control code was developed in MATLAB and utilizes proportional gain along with
Euler Parameters to determine appropriate control moments to get the spacecraft to the desired
orientation and rotation rate. The code for the ACDS system is found in the appendix of this
report. Many simulations were run in MATLAB to determine the control systems ability to
stabilize the spacecraft from various orientations and spin rates. For the first simulation, Table
5.4.1 was used as the inertia matrix for the craft. This inertia matrix represents the satellite after
deployment from the P-Pod and before solar panels are deployed. Before the panels can be
deployed, to avoid damage the satellite must be detumbled. The parameters for the detumbling
simulation are shown in Table 5.4.2 and the results of the simulation are shown in Figure 5.4.1
and Figure 5.4.2.

50. 50
Table 5.4.1: Inertia Matrix for Stowed Solar Panels
.00576 Kg*m^2 0 0
0 .0279 Kg*m^2 0
0 0 .0283 Kg*m^2
Table 5.4.2: Detumble Parameters
X Y Z
Euler Angles Initial (radians) 2.5 2 -3
Euler Angles Desired (radians) 0 0 0
Angular Velocity Initial
(radians/s)
2.22 4.44 6.66
Angular Velocity Final
(radians/s)
0 0 0
Figure 5.4.1: Angular Velocity vs. Time for Stowed Solar Panels

51. 51
Figure 5.4.2: Euler Parameters vs. Time for Stowed Solar panels
As can be seen from the above figures, the attitude control code is capable of detumbling the
spacecraft from a high spin rate. In the second test of the attitude control system, the solar panels
were deployed, and the corresponding inertia matrix is shown in Table.4.3 below. Since the solar
panels are farther away from the centroid of the craft, the inertias are larger.
Table 5.4.3: Inertia Matrix for Solar Panels Deployed
.0231
Kg*m^2
0 0
0 .0294
Kg*m^2
0
0 0 .0460
Kg*m^2
Due to the larger inertias, the control code had a harder time stabilizing the spacecraft. Shown in
Table 5.4.4 are the setup parameters for the second simulation. The results are seen in Figure
5.4.3 and Figure 5.4.4.

52. 52
Table 5.4.4: Control Simulation Setup
X Y Z
Euler Angles Initial
(radians)
1.5 2.0 2.5
Euler Angles Desired
(radians)
0 0 0
Angular Velocity
Initial (radians/s)
.22 .44 .66
Angular Velocity
Final (radians/s)
0 0 0
Figure 5.4.3: Angular Velocity vs. Time for Deployed Solar Panels

53. 53
Figure 5.4.4: Euler Parameters vs. Time for Deployed Solar Panels
While the control system may take a while to achieve a steady state, it has been shown
that the torques and control scheme are sufficient to provide attitude control for the SPDR Sat.
Further work must be done to iteratively improve the control scheme to achieve more desirable
settling times.
5.4.2 Pointing Accuracy
As stated in a previous section of this report, the omnidirectional antenna allows
for the satellite to communication with the ground station from any orientation. This eliminates
the need for precise pointing accuracies in this regard. The main limiting factor in the analysis of
pointing accuracy for the spacecraft is sun-tracking. The cosine-effect dictates that a solar panel
is most efficient when sunlight is hitting it at 90 degrees, and drops off with decreasing angle.

54. 54
Table 5.4.5: Solar Panel Efficiency as a Function of Angle
Angle (degrees) Efficiency (%)
90 100
85 91.28
80 82.63
75 74.11
70 65.79
65 57.74
60 50
55 42.64
50 35.72
40 23.39
30 13.397
20 6.03
10 1.5
0 0
The team had decided that the SPDR Sat would operate best if solar panel efficiency was
over 90%, and therefore the necessary pointing accuracy was chosen to be 5 degrees. This is well
above what is available from the Attitude Control System.
5.4.3 Sun Tracking
Another concern for the ADCS system was that it could rotate the spacecraft fast enough
to track the sun through its various angle changes during the orbit. STK was used to evaluate the
solar elevation with respect to the satellite throughout its orbit. The inclination and altitude of the
orbit were set at 51 degrees and 600 km, respectively. Shown below is the solar elevation over
time.
Figure 5.4.5: Solar Angle
From this figure, it can be seen that the solar elevation changes a total of 440 degrees in three
hours. Averaging this to find the rate of change of elevation is .04 degrees per second, or 2.4
degrees per minute. To ensure that the reaction wheels will be sufficiently able to provide this
rate of rotation, equation 1 was used.
𝑇 = 𝐽𝜔 5.4− 1

55. 55
For the value of J, .033 kg*m^2 was used (computed from Solidworks model) and for 𝜔 =
6*10^-4 rad/s was used which corresponded to the rate previously found in degrees. It was found
that the torque required to achieve this rate was 2*10^-5 N*m. This value is 100 times less than
the torque provided by our attitude control system so therefore the rate requirement is satisfied.
5.5 Primary Computing Systems Analysis
The main feature of the CPU design is that all of the components work with the CPU and
are compatible. The CubeComputer has I2C, UART, and CAN interfaces. It also comes with two
1MB SRAM for data storage and 4MB of flash storage code, as well as a microSD slot for
additional memory storage. It is found that these values provide sufficient data and memory
storage to operate an ACS code. In table 5.5.1 below there is a list of the components and their
command interface and Power Requirements.
Table 5.5.1: Component Interface
Component Interface Power Req.
Vasik Battery I2C N/A
CubeSense I2C 3.3V
CubeControl I2C 3.3V
Duplex Transceiver I2C N/A
Net Capture Device (still in
development)
I2C N/A
ESTCube-1 Camera UART 3.3V
Vacco Micro Prop Sys I2C N/A
This table shows that the CubeComputer will be able to interface with all other components at
least for the purpose of communication and command. Using incoming attitude information from
the CubeSense’s sun and nadir sensors with the CubeControl’s momentum wheels and torque
tubes the SPDR sat will be able to perform any attitude maneuver within a few degrees of
pointing accuracy. The CubeComputer will then be able to use the combination of the Vacco
micro propulsion system and the information from the transceiver to perform a burn to reach the
target orbit.
After rendezvous with the target debris, the SPDR sat will use the ESTCube-1 camera to
locate the debris at a distance of <10m. Using this information the SPDR sat will reorient itself
so that it is facing the debris and then launch the capture device. After a capture has been
confirmed the drag sail will deploy and deorbit will begin.

56. 56
Figure 5.5.1 below shows tether and end-mass image tests that were taken with Sun
equivalent irradiance. These images go from a distance of 1m away at number 1 and 10m away
at number 4 (Kuuste). While it has yet to be determined whether the ESTCube-1 camera alone
will be enough to locate the debris, there are plans for further tests to be done by the team during
construction to see if additional onboard image processing is necessary.
Figure 5.5.1: Remote image processing tests(Kuuste)
5.6 Communications Systems Analysis
There are several key factors to take into account when analyzing the Communications
subsystem. It’s important to ensure that 1) the SPDR Sat components are capable of
communicating with the designated ground station, and 2) that the data transfer rate is sufficient
to communicate the necessary information throughout the mission, which may take place at
various altitudes.
To deal with the first concern, the Friis Transmission Equation, which relates the
transmission power at one end of the communication link to the power received at the other, is
given in log form as equation (5.6-1) below. Since the selected hardware operates in the same
frequency range, that is not a concern, but it remains to be proven that the sensitivities of the
transceiver and of the ground station will be sufficient to receive signals from one another at the
designated orbit altitude.
𝑃𝑟 = 𝑃𝑡 + 𝐺𝑡 + 𝐺𝑟 + 10𝑙𝑜𝑔10 (
𝜆
4𝜋𝑅
) (5.6-1)
Here Pr refers to the power received at one end, Pt is the transmission power at the other end
(both in dBm), Gt and Gr are the antenna gains at the transmitting and receiving ends,
respectively (in dBi), 𝜆 is the signal wavelength, and R is the distance between the CubeSat and
ground station, where R and 𝜆 are in the same units. The last term is sometimes referred to as
the free-space path loss (AntennaTheory). Since the antenna gains, the frequency used, the
distance the signal travels as well as the transmitting power for uplink and downlink are all
known, the Friis equation can be used to compute the strength of the signal received by the

57. 57
SPDR Sat and the ground station during uplink and downlink, respectively. This equation can
further be used to determine the range of altitudes the SPDR Sat could occupy before
communication with a ground station is no longer possible. The ground station at Kauai
Community College (KCC) in Hawaii has been identified as a compatible ground station
(pictured in Figure 5.6.1) and will be used hereafter for sample calculations.
Figure 5.6.1: Kauai Community College Ground Station (KCC)
Its operation includes UHF and VHF frequencies in the range of 430-438MHz and 144-148MHz,
respectively, with a VHF transmit power of 500W (56.99 dBm) and 14.4dBi antenna
gain(CubeSatshop). Recalling the relevant information provided about the SPDR Sat transceiver
and antenna from Section 5.5, at an orbit altitude of 600km and an elevation angle of 00
(e.g
SPDR Sat is on the horizon), the signal power received by the SPDR Sat is -86.38dBm when
KCC is transmitting, and the signal power received by KCC is -106.71dBm when the SPDR Sat
is transmitting. Given that the SPDR Sat’s transceiver has a sensitivity of -100dBm, these values
are both well within the sufficient range of signal strength to allow for effective communication.
Given that the maximum distance between the ground station and the SPDR Sat will occur when
the SPDR Sat is on the horizon (elevation angle of 00
), the Friis equation can be used to calculate
an upper bound for the highest altitude orbit the SPDR Sat can occupy and still successfully
transmit to KCC. The signal received at the SPDR Sat has a strength of -100dBm on the horizon
at an orbit altitude of 8807.7km, which is far beyond the requirement for this mission. An
alternate value of interest is the necessary transmission power of a ground station to guarantee a
signal can reach the SPDR Sat within its transceiver sensitivity range (>-100dBm) when it’s on
the horizon in a 600km orbit. Assuming the sensitivity of the ground station will be ≤-105dBm,
a transmit power of only 43.25dBm (or 21.13 Watts) is required to communicate with the SPDR
Sat in this scenario. This is valuable information to know in the event that the customer wishes
to build their own ground station, or purchase one prebuilt, which usually costs in the range of
$35000-$50000. It should also be noted that prebuilt ground stations rarely have transmit
powers below 100 Watts, which is plenty sufficient to communicate with the SPDR Sat in an

58. 58
LEO mission. Therefore, it’s safe to conclude that signal strength and ability to communicate
with a ground station operating in the same frequency range will not be an issue for the duration
of this mission. The pertinent MATLAB code used for the above calculations can be found in
Appendix D.
The next important factor to consider is the data transfer rate. An ideal data transfer rate
is included in the transceiver description, which for downlink has a maximum value of 9600 bit/s
(CubeSatshop). The maximum uplink data rate is given as 1200 bit/s by both the transceiver and
the Kauai ground station (CubeSatshop). This information paired with access times allow for a
sample calculation of how much data can be transferred per day. Using STK with a beamwidth
of 3600
(since the antenna is omnidirectional) and selecting the Kauai Community College
ground station, access times were obtained and are displayed below with July 4th
, 2017 as a
sample day.
Figure 5.6.2: Access Times on Jul 4 2017
Figure 5.6.3: Visual Representation of Access Times on Jul 4 2017
From Figure B, the total access time during a sample day amounts to 3977.923 seconds. When
multiplied by the respective data transfer rates, this translates to 38.19MB of data that can be sent
from the SPDR Sat to the KCC, and 4.77MB of data that can be sent from KCC to the SPDR Sat.
Although this is far beyond what would be required for telemetry, it’s also likely that not all of

59. 59
the access time would be used. In the worst case scenario, assume that the SPDR Sat can only
communicate with the KCC ground station during one of the five times it passes over during a
day. Further, assume that this is the minimum duration pass, which lasts only 409.276 seconds.
During this time, the SPDR Sat can transmit 3.93MB of data to the ground station, and KCC can
transmit 491.13KB of data to the SPDR Sat. Even in this case, there is more than enough data
for telemetry commands. In the absolute worst case, if the communications hardware is
operating at its peak temperature of 600
C, the data transmission rate will approximately be cut in
half (Stengel), resulting in 1.965MB of data transmitted to the ground station and 245.565KB of
information transmitted to the SPDR Sat, which is still not problematic. If in the future the
mission is expanded and the SPDR Sat is required to send pictures or other more data-costly
transmissions to the ground station, this could become an issue, but for the current mission this is
an acceptable worst case scenario that would not cause any problems.
5.7 Thermal Systems Analysis
As expressed in Section 4.7, the thermal management portion of the Concept
Development, passive heat transfer systems are more ideal than active systems for CubeSats in
LEO. They require no power, less temperature sensitive components, and typically weigh less.
By relying on passive heat transfer we should meet our mass constraints and power limitations
while successfully maintaining the temperature of the electronics and propulsion system between
0 and 40C.
It is proposed that the spacecraft interior be lined with multi-layer insulation and that the
frame be treated with a thermal coating spray. More specifically, the 2 units including the
electronics and propulsion system will be wrapped with the insulating blanket. The frame, drag
sail case, and net capture device case will be spray coated. Through the following analysis, the
equilibrium temperature on the exterior of the spacecraft will be calculated. From this a simple
conduction analysis will be performed to determine the necessary MLI thickness to maintain the
internal spacecraft temperature above 0 C but below 40 C. Additionally, once the multi-layer
insulation thickness has been determined such that it will produce an internal temperature close
to the ideal, the emissivity and absorptivity of the exterior will be manipulated. The idea is to use
the thermal coating spray as a replacement for some layers of MLI in the event that available
space within the SPDR Sat does not permit the total MLI thickness required.
SPDR Sat’s thermal system analysis will be static in nature and cover 3 thermal loading
cases experienced throughout each orbit. The heat flux values, for an altitude of 600 km, are as
follows:
Table 5.7.1: Heat fluxes and Albedo intensities at 600 km orbit
Sunlight Eclipse
Solar IR (W/m^2) 1414 0
Earth IR (W/m^2) 234 234
Albedo 0.3 0.3

60. 60
The heat flux values presented in Table 5.7.1 are annual averages which account for
natural variations. Sources of natural variation in Earth’s infrared radiation and albedo include
cloud cover and varying terrains on Earth providing different surface reflectivity.
A few assumptions were made to simplify the analysis. It was assumed that the
spacecraft’s interior components were one body which meant the spacecraft’s internal
temperature was uniform. This assumption was valid for the preliminary analysis since it placed
all components under the strictest requirement of maintaining a temperature between 0 and 40 C.
Additionally, heat dissipated by components during operation was not considered. The team
understood that this assumption was rather unrealistic, but for the sake of calculation simplicity it
was used. In order to account for error induced by this assumption, a significant margin of
intrinsic heat was accounted for to scale down the level of required insulation predicted.
From simulations in STK, durations of sunlight, penumbra and eclipse periods were
collected and used to tailor the thermal analysis to specifically SPDR Sat’s mission. It was
determined that the SPDR Sat would spend, per orbit, an average of 64.03 minutes in sunlight,
10.65 seconds in penumbra, and 29.63 minutes in eclipse. A report detailing the duration of each
lighting case for a full day can be found in Appendix B.
The 3 locations of interested along the orbit are as follows:
1. SPDR Sat between Sun and Earth
2. SPDR Sat long side facing Earth & front face facing Sun
3. SPDR Sat in eclipse with Earth blocking Sun
Case 1 represents a scenario where SPDR Sat experiences Solar IR on its front face with
Earth IR and Albedo on its back face. The only areas taken into consideration for heat flux are
those of the front and back faces of the satellite.
Case 2 represents the instance where the satellite is no longer directly between the Sun
and the Earth but is still affected by their presence. In this case, the area on which the heat fluxes
act is composed of 2 side faces (10 cm x 30 cm) and the front face (10 cm x 10 cm).
Case 3 represents the conditions to be experienced in eclipse. Since the Sun will not be
visible to the CubeSat, the Earth’s radiation is the only active heat flux. This is expected to be the
coldest region of the orbit.
Table 5.7.2 below lists the percent coverage of each surface by certain materials. These
values are important for calculating the equilibrium temperatures at each of the 3 locations.
Table 5.7.2: The size and composition of each surface on the SPDR Sat
Front Face Side Faces Back Face
Surface Area 100 cm2
300 cm2
100 cm2
6061 Al Alloy 11.8% 22.6% 11.8%
1060 Al Alloy 88.2% 25.8% 0%
Commercially Pure
Titanium (Grade 2)
0% 0% 88.2%
MLI 0% 51.6% 0%

61. 61
The thermal analysis begins with calculating the external equilibrium temperature of the
spacecraft, Te. This can be accomplished by setting the absorbed power and emitted power
equations, Equations 5.7-1 & 5.7-2 equal to each other.
𝑄 𝐴 = 𝑄𝜀 (5.7 − 1)
𝑆0 ∗∝∗ 𝐴𝑖𝑛 = 𝜀 ∗ 𝜎 ∗ 𝐴 𝑜𝑢𝑡 ∗ 𝑇𝑒
4
(5.7 − 2)
The equation is then rearranged to solve for Te.
𝑇𝑒 = √
𝐴𝑖𝑛
𝐴 𝑜𝑢𝑡
∝
𝜀
𝑆0
𝜎
4
(5.7 − 3)
From here the values needed are the ratio of the areas which receive radiation to those
which emit radiation, absorptivity, and emissivity. The average angle at which sunlight hit the
satellite was used to calculate the effective area over which radiation was acting. The calculation
for the area ratio can be found in Appendix B. The following are the values of the effective areas
for each of the 3 loading conditions of interest.
1. A1=1/7
2. A2=2.03/14
3. A3=1/14
The following equations were used to calculate the absorptivity.
∝ 𝑠𝑖𝑑𝑒 𝑓𝑎𝑐𝑒=∝𝐴𝑙 6061 ∗
(% 𝑓𝑎𝑐𝑒 𝐴𝑙 6061)
100
+∝𝐴𝑙 1060 ∗
(% 𝑓𝑎𝑐𝑒 𝐴𝑙 1060)
100
+∝ 𝑀𝐿𝐼∗
(% 𝑓𝑎𝑐𝑒 𝑀𝐿𝐼)
100
(5.7-4)
∝𝑓𝑟𝑜𝑛𝑡 𝑓𝑎𝑐𝑒=∝𝐴𝑙 6061 ∗
(% 𝑓𝑎𝑐𝑒 𝐴𝑙 6061)
100
+∝ 𝐴𝑙 6061 ∗
(% 𝑓𝑎𝑐𝑒 𝐴𝑙 6061)
100
(5.7-5)
∝ 𝑏𝑎𝑐𝑘 𝑓𝑎𝑐𝑒=∝𝐴𝑙 6061 ∗
(% 𝑓𝑎𝑐𝑒 𝐴𝑙 6061)
100
+∝ 𝑇𝑖∗
(% 𝑓𝑎𝑐𝑒 𝑇𝑖)
100
(5.7-6)
∝ = 4 ∗∝ 𝑠𝑖𝑑𝑒 𝑓𝑎𝑐𝑒 + ∝𝑓𝑟𝑜𝑛𝑡 𝑓𝑎𝑐𝑒 + ∝ 𝑏𝑎𝑐𝑘 𝑓𝑎𝑐𝑒 (5.7-7)
Similarly, these equations were used to calculate emissivity of each face.
𝜀 𝑠𝑖𝑑𝑒 𝑓𝑎𝑐𝑒 = 𝜀 𝐴𝑙 6061 ∗
(% 𝑓𝑎𝑐𝑒 𝐴𝑙 6061)
100
+ 𝜀𝐴𝑙 1060 ∗
(% 𝑓𝑎𝑐𝑒 𝐴𝑙 1060)
100
+∝ 𝑀𝐿𝐼 ∗
(% 𝑓𝑎𝑐𝑒 𝑀𝐿𝐼)
100
(5.7-8)
𝜀𝑓𝑟𝑜𝑛𝑡 𝑓𝑎𝑐𝑒 = 𝜀𝐴𝑙 6061 ∗
(% 𝑓𝑎𝑐𝑒 𝐴𝑙 6061)
100
+ 𝜀𝐴𝑙 6061 ∗
(% 𝑓𝑎𝑐𝑒 𝐴𝑙 6061)
100
(5.7-9)
𝜀 𝑏𝑎𝑐𝑘 𝑓𝑎𝑐𝑒 = 𝜀 𝐴𝑙 6061 ∗
(% 𝑓𝑎𝑐𝑒 𝐴𝑙 6061)
100
+ 𝜀 𝑇𝑖 ∗
(% 𝑓𝑎𝑐𝑒 𝑇𝑖)
100
(5.7-10)
𝜀 = 4 ∗ 𝜀𝑠𝑖𝑑𝑒 𝑓𝑎𝑐𝑒 + 𝜀𝑓𝑟𝑜𝑛𝑡 𝑓𝑎𝑐𝑒 + 𝜀 𝑏𝑎𝑐𝑘 𝑓𝑎𝑐𝑒 (5.7-11)
For cases in sunlight, a constant maximum heat flux of S0=2072.2 W/m2
was used. In
case of eclipse, a heat flux of S0=234 W/m2
was used. The Boltzman constant value 5.6 x 10-8
W/m2
K2
was used. Plugging in these values into equations 5.7-4 -> 5.7-11 and then into equation
5.7-3, the external temperature of the CubeSat was found.