1. Final Design Report for Obsolete Satellite Capture And Removal
(OSCAR)
Version 2.0
12/11/2015
Prepared By:
Jake Adzema
Alex Austin
Austin Kubiniec
Colin Lenhoff
Alexander Malin
Ryan Moriarty
Jesse Pelletier
Rensselaer Polytechnic Institute
2. Executive Summary
Space travel has recently received increased attention from the United States government
due to the high costs associated with missions. To address the high costs associated with the
industry, there has been increased pressure to reduce the size of satellites, while retaining
performance capabilities. NASA funding has recently begun to support the design and
development of CubeSats. These satellites push forward the boundaries of engineering and
science in order to meet the stringent requirements set forth for future missions.
In addition to decreasing the size of satellites, there has also been a focus on using
commercially available parts in the satellites. These commercial off the shelf (COTS) parts are
used with the intent to reduce the production costs of the CubeSats while also increasing the
reproducibility of successful satellite systems. If a satellite can be developed as a CubeSat to
solve a specific problem then, after the successful demonstration, relatively large production can
occur with very drastic reduced costs. COTS parts allow for the satellite to be very cheap
compared to ones that use individually designed and manufactured parts.
In order to direct the development of CubeSats, space agencies have encouraged designs
to concentrate on reducing the debris that is currently in orbit around Earth. The team used this
focus in researching and analyzing components to design the Obsolete Satellite Capture and
Removal CubeSat (OSCAR). This system will be deployed in low-earth orbit (LEO), rendezvous
with and capture a debris object, and then de-orbit both the CubeSat and the captured debris.
This active debris removal concept is the best CubeSat design for cleaning space.
The team performed structural, power, thermal, radiation, data and communication
analysis, in tandem with Systems Tool Kit (STK) simulations, to determine the capabilities of
OSCAR and ensure success in deorbiting target debris. The team also developed and built a
custom payload required to capture the debris with a net in order to ensure manufacturing
feasibility and space allocation. Currently the team is awaiting sponsorship to continue more
detailed analysis and begin construction of a test model.
Providing continuation of this project, the team foresees a future in which OSCAR
CubeSats could be available at a moment's notice to send up as a secondary payload, which due
to their highly autonomous nature, will be able to bring down debris objects effectively and
safely. Over time, OSCAR will be able to have a significant impact on the amount of debris in
orbit by removing the garbage in space.
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3. Table of Contents
Executive summary………………………………………………………………………………..1
Table of Contents………………………………………………………………………………….2
List of Figures……………………………………………………………………………………..5
List of Tables…………………………………………………………………...………....……....7
Terms and Abbreviations………………………………………………………………………….8
1. Introduction………………………………………………………………………………......10
2. Project Objectives………………………………...……………………………………….…12
3. Customer Requirements……………………………………………………………………...13
4. Background on Existing Technology and Technical Standards. .…………..………….........14
4.1 Cubesat Technology Summary…………………………………...……...…………..14
4.2 Existing Research in Space Debris De-orbit Devices……………………..…………15
5. System Requirements and Design Constraints…………………………………………...…...18
5.1 Mission……………………………………………………………………………….18
5.1.1 Mission Overview………………………………………………………….18
5.1.2 Mission Success……………………………………………………………19
5.2 Structures…………………………………………………………………………….20
5.3 Propulsion……………………………………………………………………………20
5.4 Attitude Control……………………...………………………………………………20
5.5 Thermal Management………………………………………………………...……...21
5.6 Power………………………………………………………………………………...21
5.7 Command and Data…………………………………………………………………..21
5.8 Telecommunications…………………………...…………………………………….22
5.9 Payload……………………………………………………………………………….22
5.9.1 Debris Sensing Device……………………………………………………..22
5.9.2 Debris Capture Device……………………………………………………..22
6. System Concept Development and Selection…………………………..……………………..23
6.1 Structures…………………………………...………………………………………..22
6.2 Propulsion……………………………………………………………………………23
6.3 Attitude Control………………………………………………...……………………25
6.4 Thermal Management……………………………………………...………………...27
6.5 Power………………………………………………………………………………...29
6.6 Command and Data…………………………………………………………………..33
6.7 Telecommunications…………………………………………………...…………….34
6.8 Payload……………………………………………………………………………….36
6.8.1 Debris Sensing Device……………………………………………………..36
6.8.2 Debris Capture Device……………………………………………………..37
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5. A-1 Command and Data Computer System……………………………………………...87
A-2 Attitude Control System…………………………………………………………….89
A-3 Telecommunications System………………………………………………………..98
A-4 Power System……………………………………………………………………...101
A-5 Propulsion System………………………………………………………………....104
Appendix B: Component List and Cost………………………………………………………...106
Appendix C: Code……………………………………………………………………………....107
C-1 Telecommunications……………………………………………………………….107
C-2 Thermal Management……………………………………………………………...108
C-3 Power……………………………………………………………………………....109
C-4 Attitude Control…………………………………………………………………....113
Appendix D: Supplementary Figures…………………………………………………………...116
Appendix E: Team Member Contributions……………………………………………………..117
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6. List of Figures
Figure 1.1: Impact From Fleck of Paint…………………………...……………………………..10
Figure 4.1: CubeSat Missions Sorted by Mission Type………...…………...………………….14
Figure 4.2: 3U CubeSat Structure with Deployable Solar Panels.…..….……...…………….….15
Figure 4.3: Concept Image of Net Grabbing a Piece of Space Debris……….…..………….…..16
Figure 4.4: Balloon Inflated to Increase the Atmospheric Drag on a Piece of Space Debris…....16
Figure 7.1: Prograde Nodal Precession Over Quarter of a Year…………………………………40
Figure 7.2: Hohmann Transfer From 800 km to 300 km Circular Orbits………………...……...40
Figure 7.3: Retrograde Nodal Precession Over Quarter of a Year………...…………………….41
Figure 7.4: Initial Approximations of Target Debris Using MPS-120 Propulsion System...……42
Figure 7.5: Coordinate System Rotation About Z Axis by Angle ……......…………………...43θ
Figure 7.6: Block Diagram of Simulated Control Algorithm…………......……………………..47
Figure 7.7: Error Quaternions with Initial Orientation Error and No Noise……………………..47
Figure 7.8: Angular Rates with Initial Orientation Error and No Noise………………...……….49
Figure 7.9: Error Quaternions with Initial Orientation Error and Significant Noise…………….49
Figure 7.10: Angular Rates with Initial Orientation Error and Significant Noise…………….....49
Figure 7.11: Error Quaternions with Initial Orientation Error and Kalman Filter…………….....50
Figure 7.12: Angular Rates with Initial Orientation Error and Kalman Filter……………..…….51
Figure 7.13: Angular Rates with Initial Orientation Error and Initial Angular Rates……….…..52
Figure 7.14: Depiction of Earth’s Shadow Regions……………......……………………..……..54
Figure 7.15: Earth’s Shadow Approximation…………………...…………………….……..…..54
Figure 7.16: Variables Used to Calculate Gamma………………...…………………....…..……55
Figure 7.17: Beta Angle………………………………………...……………………....………..56
Figure 7.18: Power Supply…………………………………...…………………………..……...57
Figure 7.19: Power Out……………………………………...……………………….…….…….58
Figure 7.20: Battery Charge……………………………….………………………………....…..59
Figure 7.21: Excess Power…………………………………………………………………….....60
Figure 7.22: Frames per Second Calculated from FPGA Research……………………………...63
Figure 7.23: Calculation of Slant Range…………………………......…………….………...…..64
Figure 7.24: STK Simulation of Ground Tracking………………………......…..………….…...65
Figure 7.25: Payload with Sensing System Deployed…………………………………….....…..65
Figure 7.26: ESA Net Entanglement Simulation………………….....……..…….……....……...67
Figure 7.27: ESA Net Entanglement Microgravity Testing………………………......……........67
Figure 7.28: Net Device CAD Model………………….…………..……...……………………..68
Figure 7.29: Full-Scale Net Device Model Isometric View…………….…………….………....69
Figure 7.30: Full-Scale Net Device Model Front View…………………………………….........70
Figure 7.31: Full-Scale Net Device Model Side View………...………………………...……....70
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7. Figure 8.1: CubeSat Subsystem Layout………………………………………………………….71
Figure 8.2: Complete CAD Rendering with Antenna and Sensing System Payload Deployed....72
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8. List of Tables
Table 6.1: Propulsion System Selection Matrix………..……………..………………………....24
Table 6.2: ACS Selection Matrix…………...…………...……………………………………….26
Table 6.3: Thermal Requirements of Satellite Components……..………….…………………...27
Table 6.4: Properties of Different Thermal Control Paints………………….…………………...28
Table 6.5: Expected Power Draw…………………………………..……………………………29
Table 6.6: Selection Matrix for Power Supply Selection…………..…………..………………..30
Table 6.7: Power Storage Selection Matrix…………………………………...………………....31
Table 6.8: EPS Selection Matrix………………………………………...……………………….32
Table 6.9: Data Processor Component Selection……………………………...……..…………..33
Table 6.10: Telecommunications Component Selection…………………………..……..……...35
Table 6.11: Debris Capture Mechanism Selection Matrix…………………..………..………….37
Table 10.1: Risks and Challenges………………………………………………………………..81
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9. Terms and Abbreviations
ACS - Attitude Control System
AGI - Analytical Graphics Inc.
B - Byte
BIPS - Billions of Instructions per second
CAD - Computer-Aided Design
CCD - Charge-Coupled Device
CHAMPS - CubeSat High-Impulse Adaptable Modular Propulsion System
cm - Centimeter
COTS - Commercial Off The Shelf
cPCI - Compact Peripheral Component Interconnect
dBic - Decibels referenced to a circularly polarized, theoretical isotropic radiator
dBm - Decibel-milliwatts
DCM - Direction Cosine Matrix
EPS - Electric Power System
ESA - European Space Agency
FOS - Factor of Safety
FOV - Field of View
FPGA - Field Programmable Gate Array
GB - Gigabyte
I2C - Inter-Integrated Circuit
IPS - Instructions per second
ISIS - Innovative Solutions in Space
KB - Kilobyte
KB/s - Kilobyte per second
Kbps - Kilobit per second
Kg - Kilogram
LEO - Low-Earth Orbit
MB - Megabyte
MIPS - Millions of Instructions per second
MLI - Multi Layer Insulation
mW - Milliwatts
MWIR - Medium Wavelength Infrared
OSCAR - Obsolete Satellite Capture and Removal
P-POD - Poly Picosat Orbital Deployer
SEFI - Single-Event Functional Interrupts
SEL - Single Event Latchup
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10. SEPP - Solar Electric Power Propulsion
SEU - Single Event Upset
SPI - Serial Peripheral Interface
STK - Systems Tool Kit (Software by AGI)
TID - Total Ionizing Dose
TRL - Technology Readiness Level
U - CubeSat Unit
UART - Universal Asynchronous Receiver/Transmitter
UHF - Ultra High Frequency
VHF - Very High Frequency
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11. 1 Introduction
The accumulation of non-operational objects in near-Earth space poses a severe threat to
space vehicles that could be compromised, if not fatally struck, by an object in orbit. Some of
these objects are large enough that they can be tracked and monitored from Earth, but these are
only objects greater than about 10cm. Most debris, however, is too small to be tracked from
Earth, making the threat of space debris very realistic. Furthermore, this threat will not be
mitigated without an active solution. This is because the rate of objects being launched into orbit
exceeds the rate at which debris naturally de-orbits, leading to a net increase in debris. The
problem could be alleviated by simply reducing the number of objects launched into orbit, but
this would greatly hinder mankind’s ability to conduct research, exploration, utilization, and
discovery of space. The objective is to mitigate the threats posed by space debris so that new
space vehicles can continue to be launched safely, thus promoting the development of
exceptional space science.
Even if no more objects are launched into orbit, the total amount of debris in orbit could
still increase solely as a function of time by a phenomenon termed the Kessler Effect. The
Kessler Effect is the process by which objects in near-Earth orbit collide with each other and
break into smaller fragments, which in turn collide with other objects. Not only does the number
of objects in orbit increase, but their average size decreases. This means that objects large
enough to be tracked could potentially be rendered untraceable, increasing the danger posed by
these objects. This danger is emphasized in the figure below, which shows a crater 4mm wide in
the windshield of a spacecraft caused by impacting a fleck of paint 0.2mm wide. Relative to the
velocity of the spacecraft, the paint fleck was traveling between 3 and 6 km/s.
Figure 1.1: Impact From Fleck of Paint (Helvajian, 2008).
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12. There are some proposed methods by which debris in space can be de-orbited, but many
of these are still in the concept development phase (Howell, 2014). This product is expected to
show significant advancements in debris de-orbit feasibility. While there is ongoing research in
space debris de-orbit techniques, this product is especially unique due to its size. The vehicle is a
CubeSat, a nano-satellite that can be easily added to a launch vehicle already scheduled for
flight. This greatly reduces the launch cost of the vehicle. Such launching is feasible because
CubeSats are made of common 10cm cubic units, known hereafter as U’s. These units have
standard dimensions, so it is fairly easy for a launch vehicle designer to account for an extra
CubeSat payload as long as the size of the CubeSat is known. The size of a CubeSat also makes
it less susceptible to being struck by debris. OSCAR is a 3U CubeSat, which is advantageous
because it is a very common configuration that has been launched numerous times. While 3U
CubeSats are common, this vehicle is unique in that most 3U CubeSats are custom vehicles.
Furthermore, except for the payload, this product is fully comprised of flight-proven hardware.
Developing a product entirely from flight-proven commercial off-the-shelf (COTS) hardware
(with the exception of the payload) would be a great advancement. While it is crucial that the
proposed product shows advancements in debris de-orbit technologies, it should be emphasized
that fundamental operating principles of the spacecraft will show promise in more demanding
and diverse missions. These principles include the spacecraft’s structural architecture, object
rendezvous methods, and simulation techniques, among others.
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13. 2 Project Objectives
The objective of this project will be to produce a preliminary design of a CubeSat based
spacecraft to demonstrate and evaluate key COTS hardware technologies for space debris
removal. This design is separated into several subsystems designed based on capability, cost,
power, mass, and space requirements, and through assessing the risks and challenges to which
the subsystems may be subjected. The group divided responsibility for one subsystem to each of
the team members based on personal preference and expertise. This report will detail the
reasoning and methodology behind the component selection for each subsystem, as well as a
system integration plan for the entire CubeSat.
The final satellite design must consider the constraints placed by the following customer
requirements:
1. Debris to de-orbit within five (5) years of intercept by the CubeSat.
2. Significant autonomous capability in vicinity of target debris object.
3. Identify hardware failure modes and upgrade requirements to support space debris
de-orbit missions.
4. Consider extension to future non-debris missions such as Near-Earth asteroid
rendezvous or other valued science missions.
The team will submit both a detailed design report for review and a presentation to
demonstrate the entire system’s feasibility. This report reflects the input from a peer review of
the group’s midterm design review. The goal will be to move forward with prototyping based off
of this design report. The team will also present ideas for furthering our concept in the future.
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14. 3 Customer Requirements
As discussed in the introduction, space debris causes a present and future threat to
manned space flights, satellites, and other space missions. There are two major types of
customers which can be identified as being interested in the removal of debris from space: the
public customer and the private customer.
The public customer can be identified as NASA, the ESA, or any other space agency.
Their efforts in the removal of space debris are motivated by the knowledge that decreasing the
hazards of space will allow for an increase in manned space missions and the further exploration
of space. They are removing space debris with the intention of not only benefiting themselves
and their future missions, but the missions of others. The public customer is the likely source of
funding for this project, as they know that its successful development will allow other countries
or companies to use the final product at a reduced cost compared to a privately developed
project.
The private customer can be said to be a non-governmental agency operating missions in
space performing tasks ranging from taking images of Earth to delivering supplies to the
International Space Station. Two main motives have been identified for a private space company
to remove debris from orbit, the first being that it endangers their property. Satellites carry fuel to
perform orbital maneuvers which puts them out of reach of space debris, but the amount of fuel
they can carry is extremely limited and the more maneuvers a satellite has to perform, the shorter
its mission life will be. The ability to remove dangerous space debris from the orbit of a satellite
allows the company to conserve more of their satellites’ resources, enabling a longer mission
life. The secondary motive as to why a private company would be interested in the project lays
mostly in the removal of space debris that it inadvertently produced. According to NASA,
accidental explosions are the primary source of long-lived space debris in LEO (NASA Standard
8714). While regulations are in place to reduce the chances of explosions happening (and debris
being produced), accidents always happen, and the company may be required to clean up the
mess they created. This project provides a good way to de-orbit nanosatellite-sized debris in a
quick manner, as both customers require that the entire system deorbits in less than 5 years.
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15. 4 Background on Existing Technology and Technical
Standards
4.1 CubeSat Technology Summary
The origin of the CubeSat can be traced back to California Polytechnic State University
and Stanford University, where in 1999 Professors Robert Twiggs and Jordi Puig-Suari
developed the original design which has now become the standard (National Reconnaissance
Office). The original CubeSat idea came about as a way to support hands-on university-level
space exploration and education with a low-cost solution to launch science experiments into
Earth orbit (Some useful information about CubeSats). Since the year 2000, over 350 CubeSat
missions have been launched. These missions have been carried out by universities, corporations,
the military, and civilians with over fifty percent ending in success or continuing operation.
Figure 4.1 shows the distribution of all CubeSat missions sorted by mission type. (CubeSat
Database). It can be seen that the large majority of missions have been carried out in the last
three years, with commercial missions becoming the largest contributor.
Figure 4.1: CubeSat Missions Sorted by Mission Type (CubeSat Database)
The CubeSat standard definition of a U is scalable, allowing for larger structures, such as
2U, 3U, or 6U, which consist of a group of cubes attached together. The expansion of the
CubeSat industry has resulted in a large availability COTS hardware. This includes CubeSat
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16. structures of varying sizes as well as propulsion systems, attitude control systems, solar panels,
and communications systems. The prevalence of COTS hardware allows for a high degree of
customization for different CubeSat missions while keeping the total cost of the spacecraft low
(Some useful information about CubeSats). Figure 4.2 shows an example of a 3U CubeSat
structure with deployable solar panels built using COTS parts.
Figure 4.2: 3U CubeSat Structure with Deployable Solar Panels (NASA, 2015)
In addition to the CubeSat standards on the size of the structure, the compatibility with
the launch device is highly critical. CubeSats are generally launched from a P-Pod, which is a
device that is mounted within the upper stage of the launch vehicle. The P-Pod can hold
anywhere from three to six CubeSats. Once the primary payload is free of the launch vehicle, a
radiant heater severs a Vectran line to open the P-Pod door. The CubeSats are then ejected by a
spring mechanism (Isaac Nason, 2002). The P-Pod structure is usually added to a launch vehicle
designed to carry a much larger payload to orbit. In this way, the additional resources to bring
the CubeSat to space are incredibly low when compared to a dedicated launch of a large satellite.
This makes launches incredibly affordable.
4.2 Existing Research in Space Debris De-orbit Devices
There are a number of proposed methods for de-orbiting space debris which can be used
as inspiration for new ideas to be implemented on the CubeSat structure. One of these ideas,
currently being developed by the European Space Agency (ESA), is a satellite which could be
used to deploy a large net that would entangle itself around a piece of debris. Figure 4.3 shows a
concept image of what this system might look like. The satellite would then de-orbit itself,
pulling the debris along with it, where both structures would burn-up during atmospheric
re-entry. Scale testing of the net deployment device in zero gravity has been conducted to
validate launch and entangling methods around objects of arbitrary shape (Want To Snag A
Satellite).
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17. Figure 4.3: Concept Image of Net Grabbing a Piece of Space Debris (Want To Snag A Satellite)
Another de-orbit concept which could be applied to CubeSats is a device that would
rendezvous and attach to a piece of debris before inflating a large balloon. This would increase
the atmospheric drag on the spacecraft, accelerating the natural orbital decay that would
eventually cause the debris to burn up on reentry into the atmosphere. This balloon would be
made of a very thin material so that, in the event that another piece of debris strikes it, that debris
will not break up into many smaller pieces. Figure 4.4 shows a concept image of this type of
device in operation after inflating the large balloon (Global Aerospace Corporation).
Figure 4.4: Balloon Inflated to Increase the Atmospheric Drag on a Piece of Space Debris
(Global Aerospace Corporation)
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18. Still another method of de-orbiting space debris that is currently being studied is an
electrodynamic tether. This device consists of a 300 meter or longer wire that will be attached to
a piece of debris. As the wire moves through the Earth’s magnetic field, it will generate an
electric current that will slow the debris and pull it into a lower orbit. This will continue until the
debris burns up during atmospheric reentry (Parabolic Arc, 2014). It is possible that a similar
device could be adapted to be deployed from a CubeSat for smaller pieces of debris.
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19. 5 System Requirements and Design Constraints
5.1 Mission
5.1.1 Mission Overview
OSCAR’s mission can be laid out into 7 stages: launch, deployment, initialization,
rendezvous, localization, capture, and deorbit. This process takes the satellite from the ground
into orbit, where it will rendezvous and capture a piece of debris, before decreasing its altitude
until both the CubeSat and debris object burn up upon reentry into the atmosphere.
The first stage, launch, starts with the CubeSat being sent into orbit as a secondary
payload on a rocket within a P-POD. This rocket will most likely be in a sun-synchronous orbit,
at an inclination of 95-105° and an altitude between 600 to 800 km. Since this type of launch is
the most common for all observation objects, we can assume that OSCAR will most likely start
in an orbit within this range of parameters.
The next stage, deployment, starts when OSCAR is jettisoned from the P-POD. Due to
irregularities in the release mechanism of the P-Pod, OSCAR will immediately start tumbling
through space. Once it has traveled a safe distance away from its launch vehicle, the antenna will
deploy, power will turn on, and a systems check will be run on all subsystems to make sure they
have survived launch and are still operational. Then, OSCAR will begin broadcasting its status
and wait to make contact with the ground station.
The beginning of the initialization stage can be marked as the point when OSCAR makes
contact with the ground station for the first time. At this point, the attitude control system (ACS)
is activated and OSCAR begins detumbling. Once detumbling is complete, the sun is located,
and OSCAR orients itself so the solar panels can begin producing power. Finally, the ACS spins
up the craft for added stability. If all systems are listed as go, OSCAR will move on to the next
step, rendezvous. Otherwise, the mission is declared a failure and it moves to the final step,
deorbit, in order to eliminate the possibility of introducing any new space debris.
During rendezvous, the longest stage of the mission so far, OSCAR receives the orbital
parameters of a debris object close to it, and begins the process of matching its orbit through a
series of orbital maneuvers such as precession, inclination changes, and phasing maneuvers.
During this stage, all remaining subsystems are activated and start continuously operating to
keep the spacecraft powered, stable, at an acceptable temperature, and in communication with
the ground. The rendezvous section of the mission may take multiple months, or even a year,
depending on the location of the debris object, as it will be important to conserve fuel for the
deorbit phase of the mission.
Once OSCAR is about 10 meters away from the target, the localization stage begins. At
this point, the ground station accuracy is reached, and OSCAR must switch to tracking the object
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20. by deploying its stereo vision sensing system. Once the target is located, OSCAR will slowly
move towards it, stop (relative to the object, of course), and evaluate the target. The onboard
computer will be used to perform an image processing algorithm that will render a three-
dimensional image for further analysis. OSCAR will move on to the next stage of the mission if
the target meets all of the following criteria: approximately 10x10x10 cm, 2.5 kg, limited or no
tumbling, solid, sharp edges. If not, a different target is selected, and the rendezvous stage begins
again.
Once the target passes all criteria, OSCAR moves on to the capture stage, initiated by the
computer. First, the spacecraft repositions and reorients itself by moving past the object, then
completely turns around so that a retrograde burn will maneuver the satellite towards the debris.
Next, the net is fired, which will entangle the object completely. Finally, the object is pulled back
to the CubeSat to prevent a collision during the deorbit phase that could potentially create more
debris.
Finally, OSCAR moves onto its final stage, deorbit. Since the CubeSat is now facing in
the opposite direction, firing the propulsion system results in a retrograde burn which will lower
the speed and altitude of the CubeSat and attached debris object. If the maximum sized object is
captured, this final burn will bring the system to a minimum altitude of 300 km, at which drag
forces are sufficient enough to deorbit the system in less than a year. If there is any remaining
propellant, this maneuver will bring the system to a lower altitude, resulting in a faster deorbit
time. Upon reaching an altitude of 100 km both OSCAR and the captured debris object will
re-enter the atmosphere and burn up.
5.1.2 Mission Success
OSCAR can be said to have a successful mission if the following criterion are met:
1. A significant piece of orbital debris is removed from orbit
a. A significant piece would be approximately 2.5 kg and 1U in size
2. OSCAR does not contribute to the current orbital debris problem
a. OSCAR is a closed system, nothing is left in orbit
b. Critical Mission point - do not want to contribute to the problem
3. The system completes its mission in under 5 years
The success of the mission can be evaluated following the de-orbit of OSCAR. Any
mission failure should result in a full revaluation of OSCAR, as the last thing this project should
do is contribute to the problem at hand.
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21. 5.2 Structures
Launch is a very intense process and the phase in which structural failure is most likely to
occur, so it is most important to have a structure that will survive without exception. Of course,
the structure is useless if it cannot support all the mission systems. While there is no standard
electronic form factor for CubeSats defined by rules, the PCI-104 form has become standard
through practice and the structure should support this. Despite the lack of electrical form factor
requirements, there are several constraints that are imposed on the structure by the CubeSat
Design Specification, mostly regarding materials, rail geometry, and size. To conform with the
specification, the structure must be made of aluminum for compatibility with the P-POD
deployer. Furthermore, the exact size specifications must be met and a valid size of CubeSat
must be chosen. These valid sizes range from 1U to 3U+.
5.3 Propulsion
The propulsion system needs to be able to deliver OSCAR to the target through orbital
maneuvers, rendezvous with the target in close proximity, and then reduce the periapsis of the
orbit. The target debris will be focused in the 600-800 km orbit range at an inclination between
95° and 105°. After rendezvousing with the target debris, the propulsion system will need to be
able to reduce the perigee of the two connected bodies down to at least 300 km to ensure deorbit
within approximately one year. Extra fuel can be used to lower the perigee even further in order
allowing for deorbit within a shorter timeframe, but this is heavily dependent on the fuel used to
rendezvous with the debris. While performing these tasks, the propulsion system needs to
operate within the power budget, temperature limits and structural/vibrational limits imposed by
the mission. The propulsion demands stated above are anticipated to readily be met by Aerojet
Rocketdyne’s MPS-130 CubeSat propulsion system, which is currently under development.
5.4 Attitude Control
The attitude control system (ACS) must provide adequate pointing accuracy to orient
OSCAR towards any desired heading. This may be at an angle for optimal solar array
orientation, nadir pointing, or orienting the capture mechanism towards a debris target. To
accomplish this, the ACS must provide orientation control about all three body axes.
Furthermore, the ACS should ideally have redundant mechanisms in case of component failure.
Finally, The ACS must receive data from sun sensors and stereo cameras to perform rendezvous
maneuvers or localization maneuvers. Hence, the ACS must be able to interface with these
external sensors over standard interfaces.
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22. 5.5 Thermal Management
The thermal management system is required to keep all components of OSCAR within
their accepted temperature ranges. This ensures that all parts of the satellite are able to operate at
maximum efficiency and maximum lifespan. While OSCAR is in direct view of the sun, internal
temperatures will rise. This extra heat must be radiated out of the spacecraft. While the
spacecraft is in eclipse, the internal temperature will drop significantly. The spacecraft must
produce enough heat in this period to keep the internal temperature within the acceptable
operating range of all of the components. In addition, the spacecraft must be able to protect itself
from harmful radiation; this can be done by using a reflective coating on the surface of the
spacecraft as well as using radiation tolerant components.
5.6 Power
Only rigid body-mount solar panels were required to meet the power source needs for the
craft. At our estimated orbital altitude and inclination, continuous power consumption never
exceeds or nears the power supplied by a single solar array in a suboptimal position. No electric
propulsion is featured on the craft and there are no power-hungry scientific payloads to present a
large continuous power draw. A multi-channel EPS is capable of supplying peak power
demands, including the propulsion start-up power which represents the largest possible power
draw. Such undemanding power requirements allow for the use of a small battery, which takes
up relatively little space in the system.
5.7 Command and Data
Autonomy of the craft can be accomplished through the on-board computer which
features more than enough computing power to calculate its necessary maneuvers when given
correct target data, as the size of these files is on the order of tens of kilobytes (KB), and the data
transfer rates are on the order of tens of KB/s. Furthermore, it can coordinate the control of all of
the satellite's peripherals together and store gathered data for later transmission or use. Besides
the camera, these tasks require relatively little processing power and memory, so the remaining
resources will be allocated to maximize the usefulness of the camera. The parameters governing
the performance of the camera are its frame rate, and its resolution. Because the resolution is
fixed at 2594 x 1944, the remaining processing power will be used to maximize frame rate.
21
23. 5.8 Telecommunications
The telecommunications system is required to transmit and receive data between the
CubeSat and ground station. This data may include status updates, debris location information,
and commands to move forward with the next step of the mission. The combination of a
transceiver and antenna must have adequate transmission power to reach the ground station and a
high enough sensitivity to receive ground signals. As these criteria are governed by the distance
between the CubeSat and ground station, they must be satisfied at any altitude during the
mission. Finally, as it is assumed that access to a ground station will be limited to once per day
for one pass of the spacecraft, the telecommunications system must be capable of data
transmission rates adequate enough to transmit and receive all relevant mission data in a short
amount of time.
5.9 Payload
5.9.1 Debris Sensing Device
Difficult tasks, which have not been entirely considered in CubeSats, present themselves
in the topic of sensing. Ground sensing systems can aid the CubeSat in nearing its target debris,
but these systems have limited resolution for small debris. This limitation demands a sensing
system for the satellite which can detect its debris from around 10 meters away, providing
information of not just direction, but also distance, of the object from OSCAR. Absolute mission
success also requires a system which can provide insight concerning the target itself, not all
debris is possible or safe for our craft to remove. For example, the ESA estimates that over
60,000 pieces of liquid space debris in the form of NaK liquid-metal coolant exist, (Mehrholz)
and OSCAR’s sensing system needs to be able to determine if the debris it has come across is
liquid, or any other sort of danger.
5.9.2 Debris Capture Device
It is recognized that there is not a commercially available object capture device available
at this time for integration within a CubeSat. With that in mind, the project team has decided to
design a custom system for OSCAR. The device must fit entirely within the volume of 1U. In
addition, it must be capable of capturing at minimum a 10x10x10 cm piece of debris with a mass
of 2.5 kg. This debris object will be of arbitrary shape and may have some degree of tumble. In
carrying out the design, reliability will be favored as it is not expected to be feasible to have a
redundant capture device on board the satellite. For this reason, the capture system must be
highly reliable, as failure of this part would result in an inability to capture a debris object and
thus failure of the mission.
22
24. 6 System Concept Development and Selection
6.1 Structures
Structural determinations were made largely on factors brought about from other
system’s needs. Sheer size requirements of the large 1U propulsions system and multiple other
bus systems required 2U of space alone without accounting for the payload. For complete
compatibility with the CubeSat Design Specification, a CubeSat must be 3U+ or smaller and so a
3U structure was chosen, giving the payload the option to occupy the bonus volume if necessary.
Few COTS options exist for 3U structures, only two that fully met the specification
standards were found. These two structures were very similar, choosing between the two
required consideration of the final design’s details, before the final design existed. Mass
differences between the two structures differed only by a few grams, while both were structurally
proven through many launches. Selection was decided when the communication subsystem
determined its choice of antenna. The antenna required mounting using a standard front or rear
solar panel mounting pattern, which the 3U structure by Pumpkin Inc. could only accommodate
on one of the CubeSat’s ends. Unfortunately, this antenna would block the propulsion system’s
exhaust on one end or the payload subsystem’s operation on the other, but the 3U structure
offered by Innovative Solutions in Space presents solar panel mounting holes in two locations in
the middle of the CubeSat. This feature is useless for actually mounting solar panels, but is very
useful for placing the antenna, which uses the solar panel mounting pattern, in the middle of the
spacecraft. Proper accommodation of the best candidate components can only be fulfilled by the
3U structure designed by Innovative Solutions in Space, and is the primary influencing factor in
its incorporation.
6.2 Propulsion
Multiple systems were researched during the selection process for the propulsion system
which can be seen in the below table. Table 6.1 below has 5 of the 15 systems that were looked
at, representing the most probable system of each selection type (Appendix D). Scores were
determined to be from 0 to 3 for each metric with zero representing the worst and three being the
best. A scaling factor of 0.3 biases the scores to emphasize the total impulse and a factor of 0.1
reduces the importance of max thrust. Mass was not apart of the selection matrix due to most
data sheets excluding that characteristic and the assumption that propulsion systems would be
using most, if not all, of the mass allowed for the size as per CubeSat specifications.
23
25. Propulsion System Metric Selection Scores
Metric Busek
BHT-200
Busek
BIT-1
Aerojet
Rocketdyne
MPS-110
Aerojet
Rocketdyne
MPS-130
Aerojet
Rocketdyne
MPS-160
Type Scaling
Factor
Hall
Effect
Ion
Thruster
Cold Gas CHAMPS SEPP
Volume .2 2 0 3 2 1
Power Usage .2 0 1 2 2 3
Total Impulse .3 3 3 0 1 3
TRL (Technology
Readiness Level)
.2 1 1 3 2 0
Max Thrust .1 0 0 1 3 0
Total 1.5 1.3 1.7 1.8 1.7
Table 6.1: Propulsion System Selection Matrix
Based on the selection criteria outlined in the above table, the MPS-130, manufactured by
Aerojet Rocketdyne (Cubesat Modular), was chosen as the propulsion system for use on the
satellite as it received the highest score out of the systems that were found. This system should
meet the mission criteria in the best overall way. Cold gas systems were omitted mainly due to
the limit in total impulse that could be provided. Ion thrusters’ and hall effect thrusters’ main
limitations are the power draw that they require and the low TRL. SEPP was an integrated
system that included a solar panel array, but the TRL was that of a concept stage. The CHAMPS
system was the best system to achieve the orbit maneuvers that are necessary for the mission
within power capabilities.
The main disadvantage of the MPS-130 propulsion system is that the new propellant does
not have enough data and testing performed on it. However, if this becomes a hindrance in future
production, the MPS-130 can be replaced by the MPS-120 offering a flexibility in producibility
at a cost of range. The MPS-120 uses a tested hydrazine propellant with MR-142 engines, but
offers an estimated 130 m/s less ∆V than the MPS-130. Currently the MPS-120 and MPS-130
product development stages are proceeding with positive results (Carpenter, 2011).
The MPS-130 system takes up 1U of volume, has a dry mass of 1.3 kg, and has a fuel
tank that can carry 0.36 kg of fuel. The MPS 130 operates with AF-M315E (low-toxicity
propellant), has an estimated ∆V of 340 m/s for a 4 kg CubeSat, and utilizes four 1N rocket
24
26. engines. Impulsive maneuvers and thrust vectoring will be possible through the use of this
propulsion system, which will have enough fuel to rendezvous with debris and then place the
debris in a low-perigee orbit to ensure deorbit of the debris within a few years time before the
five year limit of the mission. Aerojet Rocketdyne has not yet provided a cost for either
propulsion system.
6.3 Attitude Control
Reaction wheels, momentum wheels, magnetorquers, and propellant were all considered
as possible attitude control mechanisms for OSCAR. Momentum wheels are generally very
massive and very complex, and were consequently removed from consideration. Propellant was
also quickly abandoned because OSCAR relies on its propulsion system to de-orbit. Requiring
the propulsion system to also serve as the main control actuator would siphon fuel needed to
enter a low-perigee orbit to ensure deorbit. The solar panel array has magnetorquers built into
their infrastructure, so it was determined that these magnetorquers would provide redundancy to
a separate, primary actuator system. Thus, the concept selection matrix below shows only
reaction wheel systems and their specifications.
25
27. Attitude Control System & Metric Scores
Component
Name
iADCS-100 BCT Micro
Reaction Wheel
XACT ACS BCT FleXcore
Manufacturer Berlin Space
Tech.
Blue Canyon
Tech.
Blue Canyon
Tech.
Blue Canyon
Tech.
Cost $154,000** N/A N/A N/A
Mass 345g 130g (x3) 850g 850g
Dimensions 10.39 x 9.0 x 3.1
cm
4.3 x 4.3 x 1.8
cm
10 x 10 x 5
cm
10 x 10 x 5
cm
Power 0.5V (nom) -
1.8V (max)
0.1V - 8V 0.85V - 2.83V N/A
Power Supply 5V 12V (variable
down to 8V)
12V (variable
down to 5V)
28V (variable
down to 5V)
Temp. Range -20℃ to +40℃ N/A N/A N/A
Max Speed N/A 6500 RPM 6000 RPM N/A
Max Torque .000087 Nm .004 Nm N/A N/A
Pointing
Accuracy
.0008° (.055° in
roll)
N/A .003° (.007° for
third axis)
.002°
Other
Components
3
magnetorquers,
3 gyroscopes, 1
star tracker,
nadir tracking
None N/A Magnetorquers,
2 star trackers
Table 6.2: ACS Selection Matrix (**Model with simpler software sold for $82,500)
Upon completion of the selection matrix, Berlin Space Technologies’ iADCS-100 had the
most desirable specifications in many of the categories. While the BCT Micro Reaction Wheel
seems to have the smallest mass and volume, three reaction wheels would need to be employed
to gain full control of OSCAR. The total mass and volume of three reaction wheels would then
exceed those of the iADCS-100 system. Lastly, the iADCS-100 system uses an I2C interface that
can allow for the easy integration of external sensors, such as sun sensors or star sensors for
improved pointing accuracy.
26
28. 6.4 Thermal Management
Thermal management is essential to the survival of all components of OSCAR. The average
temperature must be kept within the acceptable temperature range of all parts, listed below.
Component Temperature Range
(℃)
Radiation Tolerance
Propulsion 5 to 50 Unknown
Attitude -20 to 40 Unknown
Command -40 to 60 100 krads
Transceiver -20 to 50 Unknown
Antenna -30 to 70 Unknown
Battery -10 to 40 15 krads
EPS -40 to 85 10 krads
Sensors -10 to 55 Unknown
Table 6.3: Thermal Requirements of Satellite Components
A small patch heater is required to keep the electronics within the accepted range. A
Kapton heater, capable of operating at up to 120 ℃, will be integrated to keep temperatures
within the optimal range. This heater will be placed close to the propulsion system, which has the
highest minimum temperature. Kapton produces heaters of various voltages and watt densities;
after further thermal and power analysis, the KHLV-0502/2 was chosen, which is the smallest
heater Kapton produces. This heater is small enough to be placed next to the propulsion system,
keeping it above its minimum temperature. In addition, it is not too power hungry, so that it is
not a huge draw on the power supply. It is 1 cm by 5 cm, and produces 2 watts of power at
maximum. (Kapton® (Polyimide Film) Insulated Flexible Heaters).
The best way to keep the temperature low enough when OSCAR is in full view of the sun
is to use a thermal control paint. Because the satellite is so small, no additional cooling methods
such as heat pipes, radiators, or louvers can be installed. Instead, all of the excess heat must be
radiated directly off the satellite’s surface. This means that the emissivity of the surface must be
as high as possible. This will ensure that the maximum amount of heat is lost to outer space. In
addition, the absorptivity should be as low as possible, so that a minimum amount of heat from
the sun is absorbed by the sunlight. After researching various types of paint, AZ-70-WIZT
27
29. produced by AZ Technology was chosen due to its low absorptivity and high emissivity. One
important thing to note is that the entire surface of the spacecraft will not be coated in this paint;
most of the surface will be taken up by body solar panels, so only the remaining uncovered area
will be coated. (Spacecraft Thermal Control and Conductive Paints/Coatings* and Services
Catalog)
Paint Name Color Absorptivity Emissivity
AZ-70-WIZT White 0.1 0.91
RM-550-IB Black 0.97 0.91
AZ-3700-LSW Metallic grey 0.23 0.28
AMJ-400-IG Green 0.56 0.9
Table 6.4: Properties of Different Thermal Control Paints
Radiation protection must also be implemented. Thermal control paint is also effective
against radiation because it has a high reflectance. This means that a majority of cosmic radiation
will be reflected off the surface. In addition, due to the aluminum structure, a large amount of
radiation will be unable to penetrate the surface. Lastly, radiation tolerant components were
selected for all systems, to ensure the longevity and accuracy of the mission. Although it is
impossible to protect our satellite from 100% of radiation, the steps taken should greatly reduce
the amount of radiation that could potentially harm the electrical components of the satellite.
28
30. 6.5 Power
The power subsystem is essential in ensuring that all of the other subsystems have the
amount of electricity needed for the entire duration of the mission. Table 6.6 shows the current
minimum and maximum power draws for each of the subsystems. This data acted as a guide for
the component selections made for the power subsystem.
Power Requirements
Max (W) Min(W)
-Telecom-
Antenna 2 0.02
Transceiver 1.7 0.2
-Propulsion-
Engine 5 0
-Command-
Computer 1.5 1.5
-Power-
Battery Heater 0.3 0.1
EPS 0.4 0.4
-Attitude
Control-
ACS 1.8 0.5
-Thermals-
2 0
Total 14.7 2.72
Table 6.5: Expected Power Draw
The power system is divided into 3 components: power supply, power storage, and power
management. The selection matrices for each of these components is detailed below.
29
31. Power supply is defined as the component responsible for generating or collecting the
power to be stored or used by other subsystems. For this project, the power supply device was
chosen. Other options, such as nuclear power or electromagnetic tether, were also investigated,
but thrown away due to their irrelevance to our project.
Power Supply Metric Selection Scores
Metric Clyde Space ISIS Clyde Space GOM Space
Name SP-L-S3U-0
016-CS
3U Solar
Panel
SP-L-S1U-000
2-CS
P110
Area (U) 3 x 1 3 x 1 1 x 1 1 x 1
BOL Power (W) 8.63 6.9 2.46 2.3
BOL Vmpp ( V) 16.45 2.3 4.70 4.6
Mass (g) 135 150 42 29
Total 3 0 2 1
Table 6.6: Selection Matrix for Power Supply Selection
When the scoring was done on the Table 6.7 above, the components which were most
desirable to the project needs were highlighted in bold. The initial selections were done with both
1U panels and 3U panels to see if there was any benefit in selecting three 1U panels instead of
one 3U panel. The total highlighted in bold was chosen as the superior solar panel for the 1U or
3U category. Following the selections, the two schemes were compared using the following
equations:
3U: =62.93 W/kg135g
8.63 W
1U: =58.57 W/kg42g
2.46 W
Where it can quickly be seen that the power density, or watts generated per kilogram, is
higher in the 3U solar panel than in the 1U solar panel. After reviewing this data, the Clyde
Space 3U solar panel was chosen for the design.
The Power Storage component component is required to store the necessary power to
keep electric components functioning through eclipse and augment the energy supplied by the
30
32. solar panels in any power-intensive maneuvers. The battery charge needed to make it through a
35 minute eclipse is estimated to be 1.59 Wh, based off of the minimum power needed in Table
6.6. Using this number (plus a factor of safety of 1.5) and an additional amount for orbital
maneuvers, the minimum power that would need to be stored is 2.385 Wh. Using this number,
Table 6.8 was constructed:
Power Storage Metric Selection Scores
Metric Clyde Space Clyde
Space
Clyde Space GOM Space GOM Space
Name 10Wh 20Wh 30Wh NanoPower
2P-2S
NanoPower
BPx6 Cell
Size(mm x mm
x mm)
90.2 x 95.8
x 9.95
90.2 x
95.8 x
15.75
90.2 x 95.8
x20.4
93.39 x
87.44
x 22.86
91.68 x 85.9
x 40.00
Mass (g) 156 253 350 240 370
Discharge
Voltage (V)
6.2 6.2 6.2 8.4 6-10.8
Capacity
(WHR)
10 20 30 38.5 21.8
Rank 1 2 3 5 4
Table 6.7: Power Storage Selection Matrix
In the above table, a bold value corresponds to it being best for that metric. Due to its
minimal size, low mass, and ability to meet the project's power needs, the Clyde Space 10Whr
battery was selected as the CubeSat battery. Running the CubeSat in eclipse with the minimum
load and full battery gives 220 minutes of power, and operating with a full power draw will
provide 40 minutes of power. Proper power management will allow the satellite to perform all
necessary maneuvers with very few limitations on the operations which can be performed in
eclipse.
The power management component of this subsystem is fulfilled by the Electric Power
System(EPS). The EPS is necessary to prevent overcharging the battery and to ensure a proper
power regulation to and from all of the other subsystems.
31
33. EPS Metric Selection Scores
Metric Clyde Space CubeSat Kit Crystalspace
Name 3rd Gen. EPS Linear EPS P1U
Height(mm) 15.24 13 7
Mass (g) 86 58 80
Number of Outputs 10 6 6
Rank 1 2 3
Table 6.8: EPS Selection Matrix
The while it underperformed the other 2 EPSs slightly in the weight and height
categories, the Clyde Space EPS was chosen as it offered a greater number of outputs compared
to the two other systems, allowing for more wiring flexibility or the addition of science
components in future missions. The Clyde Space EPS also has specific integration specifications
with the Clyde Space Solar Panels and Batteries, which were previously selected for the product.
32
34. 6.6 Command and Data
Board Power
(in Watts)
Processing
(in Hertz and
Instructions
Per Second)
Memory
(in Bytes)
Interfaces Radiation
Intrepid System
Board
300mW 400 MHz
512MB NAND
flash MicroSD Unspecified
NanoMind
A3200
132mW 66 MHz
128MB NOR
flash
I2C, UART,
CAN-Bus
Unspecified
NanoMind
A712D
232mW 8-40 MHz
2x16MB NOR
flash CAN-Bus, I2C Unspecified
Proton200k Lite 1500mW
Floating
Point:
1000 MHz
1200 MIPS
8GB
flash
I2C, UART, cPCI
SEL: > 63 LET
SEU: < 1per 1000-day
TID: 100 krad (Si)
SEFI:
100% recoverable
CubeComputer <200mW 4-48 MHz
2GB
MicroSD
I2C, UART, SPI
SEL: Mentioned, not
given
SEU: Mentioned, not
given
Table 6.9: Data Processor Component Selection
Table 6.9 compiles the findings of the group’s search for a component to satisfy the
Command and Data subsystem. The search returned only two results that were
radiation-hardened and, therefore, space-capable for a 5-year mission. These components include
the Proton200k Lite (Space Micro) and CubeComputer. The CubeComputer datasheet makes
claims of “SEU protection [... and] SEL protection” but includes no data on which to base these
statements (Cube Computer). The team attempted to reach out to the manufacturer through the
cubesatshop messaging system, but met with no success. Based on this analysis, the team chose
the Proton200k Lite, the only CubeSat-compatible component that can back up its
radiation-proofing, to form the Command and Data subsystem.
Other radiation-hardened boards from the cubesatshop reference were also considered,
such as the Andrews models 160 and 110, the Q-stack, and the Q6 (CubeSatShop), but like the
CubeComputer before them, they also neglected to provide meaningful radiation data. The
33
35. Proton200k Lite expects <1 SEU (Single Event Upset) per 1000 days, which is shorter than the
timeline of the satellite's operation, so additional precautions were taken to prevent state error.
Because the Proton200k Lite was the only component to meet the design standards and
specifications set forth by the team with regards to the mission design, there was no need for a
weighted decision matrix. For the sake of argument, the team would’ve prioritized low power
and low cost above all else. This is because the satellite requires very little processing power and
memory to execute all functions aside from the camera sensors, which requires as much
processing power as we design based on its frame-rate.
6.7 Telecommunications
The telecommunications system consists of two main components, the antenna and the
transceiver. The first thing that was considered when selecting these components was the
transmission frequency range that the system would be operating in. Two ranges were
considered, Ultra High Frequency (UHF), which operates at about 100 to 500 MHz or S-Band,
which operates at about 2 to 4 GHz. Both of these ranges are below the 5 GHz threshold when
transmission issues due to poor weather can occur due to attenuation as the signal passes through
the atmosphere (Anderson).
S-Band transceivers bring the distinct advantage of increased data transmission rates over
a UHF transceiver. However, they often require an antenna with some degree of pointing
accuracy and are more expensive. As OSCAR’s mission is not purely scientific but rather to
perform the specific task of removing a piece of debris, there is no need to transmit large
amounts of data, such as high resolution images down to Earth. For this reason, the UHF
frequency range was chosen. This will also allow the CubeSat to employ an omnidirectional
antenna that will eliminate any ACS requirement for transmission pointing accuracy.
34
36. Table 6.10 summarizes the two most promising components that were compared for this
project.
Telecommunications Components Summary
Antenna Transceiver
Name
UHF Turnstile
Antenna
ISIS
Antenna
System
ISIS
Transceiver
NanoCom
AX100
Maximum Power
Consumption 10 W 2 W 1.7 W 2.64 W
Mass 30 g < 100 g 85 g 24.5 g
Communications
Interface I2C I2C I2C I2C
Uplink/Downlink Data
Transfer Rate N/A N/A
1.2 kbps/9.6
kbps 0.1-115.2 kbps
Table 6.10: Telecommunications Component Selection
Ultimately, it was decided to choose the Innovative Solutions in Space (ISIS) transceiver
(ISIS VHF) over the NanoCom AX 100 (NanoCom Communication Modules). A distinct
advantage of the ISIS transceiver is that it supports a full duplex mode. This means that it can
both transmit and receive information at the same time. For this project it is assumed that there
will be limited access to a ground station and thus this ability will ensure that the data transfer to
and from the spacecraft can be maximized in the limited time that is available. This advantage
outweighs the lower mass of the NanoCom unit.
Both the UHF Turnstile (GOMSpace) antenna and the ISIS antenna (Deployable
Antenna) consist of four monopole aerials which, when deployed, form an omni-directional
system with minimal blind spots. This will ensure that communication can occur successfully
independent of the orientation of the satellite. Both systems also feature a redundant deployment
mechanism, in the form of two heating elements per antenna, that melt a wire to release the
spring loaded antenna after orbit insertion. It must be noted, however, that this mechanism is an
additional component that must be purchased and installed for the Turnstile antenna. Ultimately,
the UHF Turnstile antenna was chosen because it is able to be mounted in the middle portion of
the structure. This will clear an opening at the top of the satellite for the protrusion of the debris
capture and sensing systems.
35
37. 6.8 Payload
6.8.1 Debris Sensing Device
A search of technologies currently being looked into for rendezvous with CubeSats
turned up very little, with most information regarding the use of small radar distance sensors
used in automobiles being converted for CubeSat use. With few other options being seriously
studied it was decided to draw on technologies used in robotics and automation to find candidate
technologies for sensing. Many were found, and each candidate technology was evaluated
individually in a qualitative overview, ultimately leading to a clear answer. Stereo vision was
ultimately selected, and brief summaries of each qualitative analysis is given here:
Radar: Small radar sensors are being looking into for use in automotive and other
consumer applications. These sensors are promising for use in space and provide excellent
distance resolution without large power consumption. However, these sensors are
one-dimensional, so their use would need to implement a method of scanning with them, either
by orienting the CubeSat or by mounting them on a electromechanical scanning device. Quickly
orienting the Cubesat with the selected reaction wheels is not possible, and constantly running
electromechanical systems pose great reliability risks in space. These two issues rendered this
method infeasible.
Laser: Laser-based sensing methods were quickly deemed infeasible as they all rely on
the assumption that the target will have optimal optical properties at the laser’s wavelength. With
the size of debris OSCAR targets, optical properties vary greatly and just about anything is
possible. Furthermore, the lasers must provide greater intensity than the sun for their incident
beam to be detectable, which is difficult. As if that wasn’t enough, stars in the background could
be confused by the laser detector as reflected laser, putting the final nail in the coffin for this
concept.
Single Camera: Cameras overcome the limitations of one dimensional sensors by being
two dimensional. Limitations seen by lasers do not affect cameras that use a wide spectrum of
wavelengths, such as normal commercially available cameras. This method offered a very
distinct advantage in that COTS hardware is available, but there were still limitations. Some
methods exist for using a single camera to determine distance, but actively evaluating the target
would still prove difficult as debris may not be static.
Multi-Camera Array (Stereo): To overcome the evaluation limitations of single cameras,
a second camera can be added to allow for stereo vision. Stereo vision can be used to create a
point cloud of what lies in view, this 3D data provides a huge advantage in evaluation of the
36
38. target, allowing for estimations of volume, mass, and observation of hazards such as sharp
corners and tumble. No COTS stereo vision system exists, but payload development is expected.
Also, a great deal of processing power is required for this system, though our computing system
was calculated to be able to meet the requirements of it. These issues do not significantly reduce
the feasibility of this method, making it the most suitable for this mission.
Structured Light or Time of Flight Cameras: While structured light cameras and time of
flight cameras are very different technologies, both suffer from similar feasibility issues as
laser-based methods as both rely on projecting light. Overpowering light from the sun makes
these systems infeasible as structured light projectors have a tough time giving out more light
than the sun.
6.8.2 Debris Capture Device
The first step in the design phase of the debris capture device was selecting the capture
method. Based on the requirement set forth in section 5.9.2 that the entire device fit within 1U,
the team selected two potential options to research which are summarized in Table 6.11 below.
Debris De-Orbit Mechanism & Metric Scores
Metric Weight Factor Net Cargo Bay
Size 0.25 0 0
Cost 0.15 3 3
Power Usage 0.15 2 2
Variety of Objects 0.25 3 1
Number of Objects 0.2 0 1
Total 1.0 1.50 1.20
Table 6.11: Debris Capture Mechanism Selection Matrix.
Each item was given a score from 0 to 3 for each metric to indicate its relative advantage
and benefits to the success of the mission. The highest scores for each metric have been bolded.
Completion of the concept selection matrix has shown that the net mechanism shows the greatest
advantage. It is expected to be very low-cost in terms of concept development, manufacturing,
and testing, and it is expected to use little power to deploy the net. Perhaps most importantly, the
net should be able to secure a very large range of objects. With the proper net design, an object
37
39. should be able to be secured regardless of its spin, shape, or material composition. The object
must be smaller than a certain maximum size, but there is no minimum size to the object given
enough net thread density. This is not true of the other mechanisms, which only work for objects
having more specific characteristics. The most significant drawback to the net is that it is not
expected to be able to capture multiple objects unless it can do so in a single launch. This, of
course, is highly unlikely.
The most promising alternative to the net, according to the selection matrix, is the cargo
bay. This method would work by having a volume within the CubeSat (the cargo bay) designated
for debris storage. Upon tracking an object, a door to the cargo bay would open and the CubeSat
would have to orient itself such that it would encapsulate the debris within the cargo bay and
close the door. The largest obstacle in using a cargo bay is that an object must fit within a certain
area. Furthermore, the velocity and spin of the object have to be very well-known to ensure
nothing is damaged by the debris inside the CubeSat.
Based on this analysis, the team opted to pursue a debris capture method using a net.
Further discussion of the final design can be found in Section 7.8.2.
38
40. 7 Design Analysis
7.1 Structure
No additional analysis was deemed to be needed for the structure at this stage of the
design after reviewing the analysis performed by ISIS.
7.2 Propulsion
The propulsion analysis is separated into three categories: orbital maneuvers, close
proximity rendezvous, and deorbit of both the debris and OSCAR. Orbital maneuvers in this
context are defined as maneuvers performed to align the orbital planes of OSCAR and the debris
that is being targeted. Close proximity rendezvous in this context include the phasing burns to
align the anomalies of the satellites. The deorbit of OSCAR and the debris is analyzed as an
impulsive retrograde burn at apogee which reduces the perigee of the now coupled objects. Due
to the openness of the problem, AGI STK Astrogator was used to analyze specific ∆V maneuver
costs.
The first orbital maneuver that was analyzed in Astrogator is a small burn made at the
ascending node of the orbit in order to change the inclination of the orbit slightly. It was
determined that a change in inclination of the orbit requires a ∆V of 13 m/s per 0.1 deg of
change. The extreme cost of this maneuver emphasizes the need for OSCAR to be placed in an
orbit that is very close to the inclination of the target debris. Moving forward in the analysis, it
will be assumed that the inclination change OSCAR can perform will be 0.3 deg change at a ∆V
cost of 39 m/s. This cost is removed from the estimated total ∆V of 340 m/s resulting in a ∆V
remaining of about 300 m/s.
The second orbital maneuver that was analyzed in STK is a nodal precession change
maneuver performed for a circular orbit at 800 km with an inclination of 95 deg. This maneuver
is composed of a small prograde burn of 38 m/s at perigee resulting in a change of the
eccentricity by .01. This causes the precession rate to be 7 deg/year less than what it would have
been without the maneuver. The simulation below in Figure 5 shows the result of this maneuver
after a quarter of a year. The blue orbit path lines are the orbits of the satellite with no maneuver,
and the green orbit paths are the orbits of the satellite after the maneuver.
39
41. Figure 7.1: Prograde Nodal Precession Over Quarter of a Year (Arctic Centered)
This maneuver can be used to adjust the difference between longitudes of ascending nodes of
two orbits over long periods of time. It is estimated that OSCAR will have approximately 260
m/s of ∆V remaining after these orbital maneuvers.
The last orbital maneuver analyzed in STK is a hohmann transfer in which OSCAR is
placed in a circular orbit greater than the target debris orbit. The largest hohmann transfer
anticipated based on our mission is a transfer from 800 km to 300 km, which requires a total ∆V
of 280 m/s. This transfer can be seen below in Figure 7.2 where the first retrograde burn occurs
at the blue-brown transition and the second retrograde burn occurs at the brown-green transition
left to orbit in a circular orbit at 300km.
Figure 7.2: Hohmann Transfer From 800 km to 300 km Circular Orbits
40
42. This would be a problem because the total running ∆V is 20 m/s over the capability of the
propulsion system, but with the proper planning a nodal precession can be performed with a
retrograde burn causing the 38 m/s of that maneuver to contribute to lowering the radius of the
orbit later performed in the Hohmann transfer. However it should be noted that the nodal
precession rate is increased by 7 deg/year in the eastward direction as seen in Figure 7.3 below.
Again the orbits are modeled for a quarter year with the original orbit in blue and the orbit after
the maneuver in green.
Figure 7.3: Retrograde Nodal Precession Over Quarter of a Year (Arctic Centered)
Now that the orbit planes of the target and OSCAR are more precisely aligned, the
propulsion system must be able to perform close proximity rendezvous. Close proximity
operations include phasing maneuvers to align the true anomalies of the two orbits and are
estimated to be small compared to the other maneuvers of the mission.
Further STK analysis of this stage of the propulsion needs to be performed in order to
determine the costs of maneuvers for the close proximity rendezvous. The analysis would need
to investigate the solutions of Lambert’s problem to align the true anomalies of the orbits. Once
within a very close distance, the Clohessy-Wiltshire equations will need to be solved in order to
perform the necessary burn maneuvers required to control the craft within sensing range of the
debris.
Deorbit maneuvers performed by OSCAR after successfully capturing the debris need
more analysis in order to directly apply to the capabilities of the mission. Due to the lack of
information regarding the specifics of AF-M315E low-toxicity propellant with the MPS-130
system, the deorbit analysis was performed using properties of the MPS-120 system (hydrazine
propellant). It should be noted that this will add a buffer to the simulation values due to the
MPS-130 having 130 m/s greater ∆V capability for a 4 kg satellite than what is offered by the
MPS-120. Using STK to solve for the changes in perigee of different target masses at increasing
41
43. orbit altitudes, the following data was generated below in Figure 7.4. Data points lying to the left
of the dotted line represent targets which can have the periapsis reduced to 300 km or below.
Reducing the object’s periapsis to this altitude is expected to deorbit the object within one year.
It should be noted that these values assume no propellant used to actually rendezvous with the
target and a full tank of propellant is used in the burn. Further information and modeling will be
needed to generate the capabilities of the MPS-130 and develop a similar plot for expected fuel
amounts that OSCAR will have at the target debris.
Figure 7.4: Initial Approximations of Target Debris Using MPS-120 Propulsion System
7.3 Attitude Control
In order to properly determine if Berlin Space Technologies’ iADCS-100 meets the ACS
requirements stated in Section 5.4, a simulation was conducted using the Simulink simulation
suite operated with a MATLAB script. The simulation was run using a discrete time step of 5Hz
and using a quaternion-based parameterization, both of which accurately represent the operating
conditions of the iADCS-100. The simulation was also designed to saturate the reaction wheel
torque at 0.087 mNm, which is the maximum torque provided by any single reaction wheel in the
iADCS-100. Along with the maximum torque, it is important to have a good estimate of
OSCAR’s inertia matrix. The inertias about each of OSCAR’s primary body axes are presented
below.
.03 Kg I .03 Kg I .006 KgIxx = 0 * m2
yy = 0 * m2
zz = 0 * m2
42
44. These three inertia values, obtained from the computer-aided design (CAD) of OSCAR, make up
the diagonal of the full 3x3 inertia tensor. In this tensor, the off-diagonal terms are very close to
zero and are approximated as equalling zero.
Before delving into OSCAR’s attitude dynamics, it is necessary to first explore two
methods by which OSCAR’s body-frame coordinates can be related to an inertial frame fixed
with respect to an Earth-based observer. It is crucial to establish this relation because any desired
orientation angle sent to the control system is given with respect to an inertial frame, which is a
reference frame that remains stationary with respect to an Earth-based observer. On the other
hand, OSCAR’s on-board gyroscopes (also called gyros) measure angular rates in its own
reference frame, called the body frame. The first relationship that relates these two frames to
each other is established using Euler angles. The theory of Euler angles states that an object can
be rotated about the three axes of a stationary (inertial) frame to achieve any orientation in the
inertial frame. In order to characterize this orientation, consider just a single rotation of the body
frame axes about an inertial frame as shown in Figure 7.5 below:
Figure 7.5: Coordinate System Rotation About Z Axis by Angle .θ
Consider the body frame to be defined as having axes [X’, Y’, Z’] and the inertial frame as
having axes [X, Y, Z]. The z-axes of each frame are pointing out of the page in the figure above.
Next, consider the body frame to initially have the same orientation as the inertial frame and then
it is rotated about its Z-axis by an angle . In order to relate the inertial frame to the body frame,θ
the rotation angle and simple trigonometric relations can be used. The following relation is
obtained:
(eq. 7.3.1)
43
45. The 3x3 matrix is known as a direction cosine matrix (DCM). When the body frame is rotated
about each of its axes independently, the total DCM relating the two frames is given by simply
multiplying the DCM obtained by each individual rotation. By rotating the body frame about the
X, Y, and Z inertial-frame axes in that order, the following total DCM is obtained:
(eq. 7.3.2)
where the subscripts represent Axes 1, 2, and 3 (X, Y, and Z in this case) and and stand forc s
the cosine and sine (respectively) of the rotation angle about the subscripted axis. This DCM will
be referenced in later calculations.
Relating angular rates across two reference frames using Euler angles results in equations
with singularities. Arriving at such a singularity can cause a failure of the ACS unit (specifically
known as gimbal lock), sabotaging the entire mission. Thus, it is imperative to use an alternative
means of relating angular rates between reference frames. This method is known as the method
of quaternions (also referred to as Euler parameters). Whereas the theory of Euler angles states
that any orientation can be achieved through a series of rotations about three principle
perpendicular axes, the theory of quaternions states that an object can achieve any arbitrary
orientation in an inertial frame by performing a single rotation about a single arbitrary axis. This
axis is defined by three quaternion parameters that define the orientation of the rotation axis with
respect to the axes of the inertial frame. The magnitude of the rotation is characterized by a
fourth parameter. These four quaternions are related to the terms in the total DCM shown earlier
by the following equations:
(eq. 7.3.3)
44
46. (eq. 7.3.4)
where represents the item in the total DCM located at row , column . Quaternions areCij i j
desirable not only because they lack singularities, but because they are more robust and
calculations can be conducted more quickly than those conducted entirely with Euler angles.
Most importantly, the iADCS-100 uses quaternions when performing calculations so that is what
has been simulated in all analysis. With this quaternion mapping, the attitude dynamics can now
be properly assessed.
The inertia and torque properties mentioned previously are crucial in simulating the
attitude dynamics of the system; in this case, OSCAR and its attitude dynamics represent the
“plant” of the whole system (refer to the block diagram in Figure D.1 of Appendix D). In order
to properly simulate OSCAR’s attitude dynamics, Euler’s equations of motion for rotation are
used:
α x IωI = ω + τ (eq. 7.3.5)
In this system of equations, is the 3x3 inertia tensor, , is a column vector ofI α = dt
dω
ω
body-frame angular rates measured by the on-board gyros, and is the total torque columnτ
vector acting on OSCAR (control torques and external disturbance torques). The system of
equations above are solved for and integrated to output the vector of body-frame angular ratesα
. Once the rates are calculated, they are used to determine the rate of change of the orientationω
quaternions, shown below:
(eq. 7.3.6)
where , , and represent the angular velocities about the X, Y, and Z axes of OSCAR’sω1 ω2 ω3
body-frame coordinate system. In this system of equations, the quaternion rate vector on the
left-hand-side represents the values at the “present time step,” whereas the quaternion vector on
the right-hand-side represents the quaternions calculated at the previous time step (keep in mind
this is a discrete-time solution method). The present-step quaternions are found by integrating the
45
47. present-step quaternion rates, which can be plugged into the right-hand-side of the equation to
find the quaternion rates at the next time step once the angular rates have been updated at the
next time step as well.
While OSCAR must achieve a desired angular orientation, it must also be able to achieve
a desired body-frame rotation. During the course of the mission, OSCAR may experience
external disturbance torques caused by phenomena such as solar pressure, since OSCAR’s center
of gravity is offset from its geometric center by roughly two centimeters. The adverse effects of
the disturbance torques can be mitigated through a process called spin-stabilization.
Spin-stabilization constantly re-orients an object so that the moment arm between the geometric
center and center of gravity always changes direction. This prevents disturbance torques from
always acting in the same direction, which would cause OSCAR to experience significant
angular acceleration over time. It is important to note that OSCAR is a small craft and solar
pressure will likely not be a major issue, but it is best to have spin-stabilization capability in case
it is.
Since OSCAR needs to have control over angular orientation and rotation, a control
scheme that takes orientation quaternion error and angular rate error into account has been
implemented in simulations. Quaternion errors are calculated by first converting the orientation
quaternions to a total DCM using Equation 7.3.4 and multiplying the resultant by the transpose
of the total DCM relating desired orientation to OSCAR’s body frame. The angular rate error is
simply the difference in the actual and desired angular rates. The actual quaternions and the
actual angular rates are calculated and directly measured, respectively, while the desired
quaternions and desired angular rates are sent to the on-board computer and then to the
iADCS-100 unit. This information is known as ephemeris data (see Figure D.1 in Appendix D).
Once the quaternion errors and angular rate errors are calculated, they are each scaled by a term
known as a gain. Multiplying errors by gains allows the total control signal to be weighted more
heavily on some errors than others. Gains can also be thought of as unit converters; for example,
an orientation error of 0.1 radians may need to be mapped to a voltage of 2V across the reaction
wheels so they provide the proper control torque to re-orient OSCAR. In this case, a quaternion
gain of 20 would be appropriate. In the control scheme used for all simulations in this
preliminary analysis, the quaternion and angular rate errors are each multiplied by a gain vector.
These two resulting signals are added to create the control signals sent to the reaction wheels, as
shown in Figure 7.6 below.
46
48. Figure 7.6: Block Diagram of Simulated Control Algorithm.
The first simulation that was conducted on OSCAR simulated desired Euler angles of [0°,
0°, 0°], initial Euler angles of [90°, -165°, 120°], desired angular rates of [0°/s, 0°/s, 0°/s], and
initial angular rates of [0°/s, 0°/s, 0°/s]. This produced the slowest settling time among initial
angles that were simulated. Quaternion gains were tuned to [5, 5, 5] while angular rate gains
were tuned to 29.*[0.8, 1.2, 1]. The results are shown in Figures 7.7 and 7.8 below, illustrating
quaternion error and angular rates over time.
Figure 7.7: Error Quaternions with Initial Orientation Error and No Noise.
47
49. Figure 7.8: Angular Rates with Initial Orientation Error and No Noise.
In performing these simulations, it was assumed that the on-board gyros were ideal and
measured no noise. That assumption is not valid for any real system. For this reason, a
subsequent simulation was conducted to simulate the same initial conditions and desired
conditions but with added noise in the gyros. This was done by taking the angular rates from
integration of Euler’s equations and adding the product of a standard deviation and a
normally-distributed random number. To model an extreme case, the standard deviation value
was chosen to be 0.05 radians per second. This is about 66% of the maximum angular rate
observed in Figure 7.8 above, which represents a very extreme case that might occur only in the
event of a collision or some other means of harsh external disturbance. The results of this
simulation are shown below in Figures 7.9 and 7.10.
48
50. Figure 7.9: Error Quaternions with Initial Orientation Error and Significant Noise.
Figure 7.10: Angular Rates with Initial Orientation Error and Significant Noise.
49
51. The settling time for the error quaternions in this case is over 15% larger than in the ideal case of
no noise measurement. Fortunately, the iADCS-100 is equipped with an on-board state estimator
known as a Kalman filter. Simply put, a Kalman filter uses a physics-based model to predict
what the angular rates should be at the next discrete time step. It then compares these values to
the values recorded by the gyros and, based on the difference between the two, applies a gain to
the measured value before sending this value to the control algorithm. If the measured value
closely resembles the predicted value (the state estimate), the value is “trusted” more heavily and
is not much affected by the Kalman gain. On the other hand, if the two values disagree
significantly, a gain will be applied to bring the Kalman filter’s output closer to the state
estimate. Results of simulating the Kalman filter are shown in Figures 7.11 and 7.12 below.
Figure 7.11: Error Quaternions with Initial Orientation Error and Kalman Filter.
50
52. Figure 7.12: Angular Rates with Initial Orientation Error and Kalman Filter.
The results shown above are the most accurate simulations of OSCAR’s performance.
Figure 7.11 above, when zoomed in, shows that every error quaternion converges to within 1°
within 70 seconds.
Next, a simulation of a detumble operation was performed. OSCAR was simulated
having initial Euler angles of [90°, -165°, 120°] and initial angular rates of [6°/s, 6°/s, -6°/s]. The
quaternion gain was set to zero so the only control effort would be to set the angular rates to
zero. The simulation includes modeling of noise and a Kalman filter, with results shown in
Figure 7.13 below.
51
53. Figure 7.13: Angular Rates with Initial Orientation Error and Initial Angular Rates.
The results indicate that OSCAR should be able to recover from a moderate tumble in less than
45 seconds under these conditions, allowing the mission to quickly take shape after deployment
from the primary launch vehicle.
7.4 Thermal Management
As mentioned before, the goal of the thermal management subsystem is to ensure that the
average internal temperature of the spacecraft is to be kept within the accepted temperature range
of all components. This range was determined to be between 5 and 40°C. The minimum
temperature is defined by the propulsion system, and the maximum temperature is defined by
both the ACS and the battery. To calculate the average internal temperature, an analysis of the
heat into and out of the satellite must be performed. Heat gained by the satellite is a combination
of: heat absorbed by the sun (Qsun), heat from the sun reflected off the earth called albedo (Qalb),
heat absorbed via infrared radiation from the earth (QIR), and heat generated by the satellite
(Qgen). Heat is only lost by the satellite in one way: through radiation into space (Qrad). According
to the first law of thermodynamics, Qout = Qin. Substituting in the sources of heat, the following
equation is generated:
(eq. 7.4.1)
52
54. Heat emitted by radiation can be defined as
(eq. 7.4.2)
Where A is the area, ε is the emissivity, 𝞂 is the Stefan-Boltzmann constant, and T is the average
internal temperature.
Because the surface of the spacecraft is not all made of the same material, and therefore has
different emissivities, a more complex form of this equation must be used.
(eq. 7.4.3)
On a similar note, heat absorbed from the sun can be defined as
(eq. 7.4.4)
Where Gs is the average solar flux, Asn is area of a specific material facing towards the sun, and
⍺n is absorptivity of a specific material
The albedo can be defined as a percentage of sunlight reflected back from the Earth’s surface.
The average is about 30%, so the heat from albedo equation becomes
(eq. 7.4.5)
Where Aen is area of a specific material facing the Earth
Heat absorbed via infrared radiation from the earth is similar, but depends on the emissivity. It is
defined as
(eq. 7.4.6)
Where GIR is the average infrared flux
Lastly, heat generated by the satellite can be calculated in terms of maximum and minimum
power used by the system. These numbers were taken from the power subsystem.
By combining all of these equations, the only unknown is internal temperature. This can
be solved for in two scenarios: minimum and maximum. The minimum temperature occurs when
the satellite is in eclipse; this means that heat absorbed by the sun and albedo are both zero. The
maximum temperature occurs in direct view of the sun; this means that all sources of heat must
be considered at their maximum values. Using a MATLAB script (appendix C-2), these
temperatures were calculated to be 10.67℃ and 38.01℃, respectively. Both of these values are
within the optimum temperature range, meaning that the spacecraft should be able to function
properly throughout its entire mission. (Friedel, Jonas, and Sean Mckibbon)
7.5 Power
When analyzing the power subsystem in Section 6.5, the battery charge needed to sustain
the CubeSat through eclipse was determined as 2.385 Wh, thus, entering eclipse with any charge
53
55. less that 2.385 Wh was can been seen as a failure during the validation process. An important
assumption made in this section is that the CubeSat is in a circular or near-circular orbit
(eccentricity ~ 0).
An important part of determining how long a satellite will be in eclipse depends on how
the shadow of the Earth is modeled. The Earth (and any other planet) has two shadow regions,
Umbra Cone and the Penumbra Region, both depicted in Figure 7.14. In the Umbra Cone, the
Earth is completely blocking the light of the sun, while in the Penumbra region, the light of the
Sun is only partial blocked (NASA Technical Paper 3547).
Figure 7.14: Depiction of Earth’s Shadow Regions
Due to the extreme distance between the Sun and the Earth, the Sun can be approximated
as a point light source, which results in the cylindrical shadow seen in Figure 7.15
Figure 7.15: Earth’s Shadow Approximation
54
56. To approximate the time that the CubeSat will spend in this eclipse the following steps
were taken:
Radius of EarthR E =
adius of orbitr = r
alf angle of eclipseγ = H
(eq. 7.5.1)
Figure 7.16: Variables Used to Calculate Gamma
Following the calculation of gamma, the time that the CubeSat spends in eclipse can be
calculated by multiplying the orbital period by the fraction of the orbit that the CubeSat is in
eclipse:
rbital PeriodT = O
raction of orbit in eclipsefe = f
eriod of EclipseTe = P
emi ajor axisa = s − m
ravitational Parameter of EarthμE = G
(eq. 7.5.2)
(eq. 7.5.3)
(eq. 7.5.4)
55
57. Using these equations, the amount of time the satellite is eclipse can be calculated to be
35 minutes. For a more complex analysis, the beta angle can be used (Orbital Mechanics for
Thermal Engineers).
Figure 7.17: Beta Angle
As depicted in Figure 7.17, the beta angle is the angle between the orbit plane and the
solar plane. This can be used to accurately predict the amount of sunlight a CubeSat will see at
different angles of inclination and different longitudes of the ascending node. However, for
OSCAR’s power analysis, the beta angle was set to zero to ensure the CubeSat could endure the
worst-case power scenarios.
Utilizing the above equations and assumptions, a MATLAB model was constructed to
ensure the power system can meet all of the needs of the spacecraft. The MATLAB code is
featured in Appendix C-3. The following MATLAB plots are an example of some of the
analyzation performed on the power system. The model has the input parameters:
0° (beta angle)β =
5° (inclination angle)i = 8
7171 km (orbital radius)r =
.1 k (rotational speed of spacecraft)ωr = 0 ˆ
sec
rads
imestep 15.1085 sec (time in step of simulation)t =
The following figures depict a snapshot of the power parameters over three orbital
periods. Each period is shown in a different color.
56
58. Figure 7.18: Power Supply
Figure 7.18 shows the power supplied by the solar panels to the CubeSat as a function of
time in the orbit. In this simulation, the satellite is rotating about its z-axis for reasons explained
in section 7.3. This causes the solar panels to be at a constantly changing angle to the sun, which
results in a sinusoidal input of power due to the cosine law of solar panels (Anderson). The
power in on the graph drops to zero when the spacecraft enters eclipse.
57
59. Figure 7.19: Power Out
In contrast , Figure 7.19 shows the power out as a function. As stated in Section 6.5, the
minimum power output is 2.76 W, and this is set as the default power out when the craft is not in
eclipse. When in eclipse, the minimum power output is set to 4.72 W; the additional 2 W of
power draw comes from running the heaters on full. As discussed in 7.4, running the heaters on
full for the entire eclipse is an unlikely scenario but the CubeSat is more than prepared to handle
that power draw if need be. The other two spikes in the power draw are examples of what would
be considered standard orbital operations; the first spike is caused by a 20 minute transmission to
earth and the second spike is caused by a 5 minute burn of the propulsion system while in
eclipse.
58
60. Figure 7.20: Battery Charge
Figure 7.20 depicts the charge of the battery as a function of time. This graph was created
taking the difference of the power in and the power out, and adding or subtracting any surplus or
deficit from the charge in the battery. Any charge over 10 Wh is not allowed as that is the
capacity of the battery.
59
61. Figure 7.21: Excess Power
Illustrated in Figure 7.21 is the excess charge that the battery currently has stored. This
value has been calculated by taking the current charge in the battery and subtracting the
minimum value of power need to make it through eclipse (calculated in Section 6.6). This figure
was constructed to ensure that the CubeSat never entered eclipse with an inadequate amount of
power. The star points on the graph represent the points at which the satellite enters eclipse, and
as long as these points remain above zero, OSCAR will successfully endure eclipse. In the
future, this graph will also allow the design team to optimize the timing of orbital operations by
getting the design point( the star marker) as close to zero as possible.
60
