Critical Design Report
Contents
Team Summary ............................................................................................................................................. 5
Introduction .................................................................................................................................................. 6
Changes Made Since PDR.............................................................................................................................. 7
Mission Criteria ............................................................................................................................................. 8
Mission Statement .................................................................................................................................... 8
Mission Requirements .............................................................................................................................. 8
Top Level Requirements ........................................................................................................................... 8
Mission Success Criteria ............................................................................................................................ 9
Critical Mission Milestones ....................................................................................................................... 9
Concept of Operations ................................................................................................................................ 10
Launch ..................................................................................................................................................... 10
Deployment & Stabilization .................................................................................................................... 11
Operations and Deorbit .......................................................................................................................... 11
Satellite Sub-system Design ........................................................................................................................ 12
Structures ................................................................................................................................................ 12
Design Selection .................................................................................................................................. 12
Interfaces ............................................................................................................................................ 14
Load Analysis....................................................................................................................................... 15
TRL ....................................................................................................................................................... 19
Risk Analysis ........................................................................................................................................ 19
Payload.................................................................................................................................................... 19
STK Models for Estimated Impacts...................................................................................................... 19
Final Design ......................................................................................................................................... 21
DRAGONS Units Deployment .............................................................................................................. 26
TRL Advancement Plan........................................................................................................................ 29
Thermal Protection ................................................................................................................................. 30
Aluminum Radiation Shield ................................................................................................................. 30
Multi-Layer Insulation ......................................................................................................................... 31
Safety and Mission Assurance............................................................................................................. 37
Risks Management.............................................................................................................................. 37
Electrical Power System .......................................................................................................................... 37
2
Design Outline ..................................................................................................................................... 38
Power Collection ................................................................................................................................. 38
Energy Storage .................................................................................................................................... 40
Power Conditioning and Distribution .................................................................................................. 40
Power Budget ...................................................................................................................................... 41
Configuration ...................................................................................................................................... 42
Risk Analysis ........................................................................................................................................ 43
Interfaces ............................................................................................................................................ 43
Orbit Attitude Determination and Control ............................................................................................. 43
Design.................................................................................................................................................. 43
Attitude Control................................................................................................................................... 48
Risk Analysis ........................................................................................................................................ 49
Technology Readiness Level ................................................................................................................ 50
Interfaces ............................................................................................................................................ 51
Command and Data Handling ................................................................................................................. 52
Hardware ............................................................................................................................................ 52
Software .............................................................................................................................................. 52
Technology Readiness Level Advancement Plan ................................................................................. 54
Communications ..................................................................................................................................... 54
Design Outline ..................................................................................................................................... 54
Technology Readiness Level ................................................................................................................ 56
Interfaces ............................................................................................................................................ 56
Safety and Mission Assurance............................................................................................................. 57
Final Product ............................................................................................................................................... 58
Operations .................................................................................................................................................. 60
Risk .............................................................................................................................................................. 61
Technology Readiness Level ....................................................................................................................... 66
Verification and Validation ......................................................................................................................... 68
Verification Matrix .................................................................................................................................. 68
System Performance Requirements Flow down Chart ........................................................................... 86
Project Management .................................................................................................................................. 88
Statement of Work ................................................................................................................................. 88
3
Team Structure ....................................................................................................................................... 89
Budget ..................................................................................................................................................... 89
Launch Services ................................................................................................................................... 89
Project Cost ......................................................................................................................................... 90
Mass .................................................................................................................................................... 93
Schedule .................................................................................................................................................. 94
References .................................................................................................................................................. 95
Appendix ..................................................................................................................................................... 97
Acronyms ................................................................................................................................................ 97
Limited Life Items List ............................................................................................................................. 98
Drawings ............................................................................................................................................... 105
EPS..................................................................................................................................................... 105
Payload.............................................................................................................................................. 107
Structures .......................................................................................................................................... 113
Communications ............................................................................................................................... 120
OADC ................................................................................................................................................. 121
Gant Chart ............................................................................................................................................. 125
4
Team Summary
Bravo Aerospace Industries
800 SW 5th Ave.
Gainesville, FL 32601
Phone:
954-665-6718
Email:
lcnianarivera@ufl.edu
Project Manager:
Chief Systems Engineer:
Ian Rivera
Radina Dikova
Members:
Kathleen Hencke
Lester Tuck
Nicholas Hoffman
Brian Lucas
Chris Stewart
Michael Koester
Lee Silvernail
Noopur Joshi
Maggie Owen
Bryan Maynard
Joseph Levesque
Chris Roda
Ryan Barows
Paula Pluchino
Fred Ouellet
Daniel Hugentugler
Nathaniel Arias
Tony Diaz
5
Introduction
Crowding of the earth’s orbital space environment poses serious risks to space missions and
creates an unpredictable environment. In 2009, a hypervelocity collision occurred between two
satellites, the Iridium 33 and the inactive Cosmos-2251. This collision incapacitated the active
Iridium 33 and created thousands of pieces of debris in the process, resulting in an even more
hostile space environment. Analysis by both NASA and outside experts indicates that more than
half of the Iridium debris will remain in orbit for at least 100 years, and much of the Cosmos
debris will remain in orbit for at least twenty to thirty years.1
While the dangers of collisions between two large objects in space such as the Iridium-Cosmos
collision may be obvious, there isn’t much knowledge regarding the hazards small space debris
(less than 10 cm in diameter) poses. Nicholas Johnson, NASA chief scientist for orbital debris,
would argue that small space debris is even more dangerous than large debris (greater than 10 cm
in diameter). “The greatest risk to space missions comes from non-trackable debris.” 2 It is clear
that small space debris poses some level of danger to spacecraft. This claim is supported by the
damage caused to a space shuttle window due to a paint fleck. The damage was substantial
enough that the window was replaced.
To prevent damage caused by micro meteoroids and orbital debris (MMOD), an understanding
of the dangers must be obtained. MMOD characterization allows for improved spacecraft design
or maneuvering strategies. Bravo Aerospace Industries (BAI) has developed a comprehensive
approach to characterizing this debris.
BAI has designed the Bravo Orbital Space Satellite (BOSS) to take in situ measurements of
small space debris. The BOSS project will introduce a low cost, yet reliable solution to
characterizing MMOD while not contributing to the problem. BOSS has the potential to
revolutionize future satellite and spacecraft design and will provide the information necessary to
formulate a solution to the space debris problem.
The BOSS project is of interest to all entities involved in the space industry. From protecting the
International Space Station to ensuring communications satellites’ health, the mission will prove
to be vital in the advancement of space technology. Since NASA possesses some of the most
expensive equipment in orbit, the BOSS mission is of particular interest to NASA. The BOSS
satellite is designed for the advancement of NASA’s technological capabilities, although the
project’s implications will impact most space entities.
BAI understands the industries’ interest in low earth orbit (LEO) missions. LEO missions are
relatively cheap and simple when compared to geostationary orbit (GEO) missions. LEO debris
characterization is also of interest since all missions, from LEO missions to Mars missions, must
logically travel through LEO. The potential for impact in LEO is also much higher than in GEO
due to a higher debris density. For this reason, the BOSS satellite will operate in a LEO
environment.
6
NASA currently lacks the technology to assess small space debris with the accuracy that BOSS
will be capable of. Impact test based models such as the Satellite Orbital Debris Characterization
Impact Test or the NASA-Kyushu University tests have provided some models for small space
debris, however these are merely attempts at duplicating unpredictable collisions which is
inherently inaccurate. Analysis of test spacecraft such as NASA’s Long Duration Exposure
Facility may provide insight into the damage a spacecraft may incur during a mission; however,
it provides little information regarding the type of energy associated with different sized debris.
Ground based systems such as the European Space Agency’s Space Situational Awareness
program or the U.S. Space Surveillance Network are hardly worth mentioning since they are
limited to only measuring debris of diameter larger than three millimeters.7
With the use of the innovative Debris Resistive Acoustic Grid Orbital Navy Sensor
(DRAGONS) payload, BOSS will be capable of characterizing space debris as small as 50 μm. 5
The DRAGONS payload combines a control and data storage subsystem with an acoustic
subsystem and a resistive grid subsystem to create an instrument designed to detect microscopic
particles in polar and sunsynchronous orbits.5 Introducing this level of technology into the space
environment is necessary and will open the door for further research and development of space
materials. It will also increase the accuracy of current space models by allowing them to account
for small space debris.
Changes Made Since PDR
1. Project Management
a. The schedule has been updated to reflect a more accurate time frame for the
project.
b. The budget has been updated to more accurately represent the cost of the project.
2. Command and Data Handling (C&DH)
a. The maximum power consumed by C&DH is now 9W (previously 5W). This is
due to the final computer board selection.
3. EPS
a. More specific power requirements numbers from other subsystems has resulted in
reduced overall power need.
b. The solar array has been scaled down by 60% due to higher efficiency and
updated power budget figures.
c. A smaller capacity battery will be used due to updated power budget figures.
d. To guarantee peak power generation, a gimbal has been added to the final solar
array design.
4. OADC
a. The two star trackers were removed from OADC. The star trackers were deemed
unnecessary as the angular velocity and angular acceleration from the IMU
provides highly accurate measurements through Kalman Filtering.
b. The de-orbit module was changed to NanoSat Terminator Tape from a sail.
5. Structures
a. To reduce mass while maintaining structural strength, each of the six side panels
supporting the rectangular structure will now consist of an aluminum-honeycomb
material as opposed to just aluminum.
7
b. It was also decided to add a shelf in the interior of the satellite in order to aid the
assembly of the internal components of the satellite (specifically, the reaction
wheels and modified IMU required by OADC).
c. The orientation of the frame has been changed to 0.6m x 0.5m x 0.6m (length,
width, height) from 0.6m x 0.6m x 0.5m.
6. TPS
a. Using a phase change device to distribute heat has been deemed unnecessary.
Mission Criteria
Mission Statement
The debris field in space is a concern for all spacecraft. The particles pose a risk for impact of
spacecraft currently in space and are a concern for future spacecraft. This impact risk also poses
a threat to the science missions being conducted by these spacecraft. Particles smaller than
millimeter in size are non-trackable and can pose a certain risk to spacecraft.8
The BOSS mission is to characterize millimeter or smaller MMOD particles in Earth’s orbit. Insitu measurements from sensors will characterize three components of impact from particles. The
three measurements include size of impact, speed of impact and direction of impact.
Mission Requirements
The BOSS mission has four mission requirements. These four requirements will ensure that
BOSS is able to meet the mission demands of the customer. It will ensure the purpose of the
mission, and the maximum cost and weight are all clearly stated and met when the satellite is
complete.
01.0: BOSS shall characterize the near-Earth debris field of particles up to one millimeter in size.
2.0: BOSS shall characterize debris field using an energy detection method.
03.0: BOSS shall a have a total cost of less than 15 million dollars.
04.0: BOSS shall contain a total mass less than 180 kg.
Top Level Requirements
The top level requirements indicate how the mission requirements will be met. Each of the eight
subsystems of the BOSS mission has a set of requirements that will ensure the mission
requirements are met.




Payload (05.0)
- The payload instruments and sensors shall provide sufficient capabilities for millimeter
size debris detection.
Bus: Command and Data Handling (C&DH) (06.1)
- The C&DH shall successfully collect and process data to be transmitted to the ground
station.
Bus: Communications (COMM ) (06.2)
- The COMM shall successfully transmit “energy data” to ground station.
Bus: Electrical Power Systems (EPS) (06.3)
- The EPS shall store and distribute adequate power necessary for operation of all
components.
8





Bus: Orbit Attitude Determination and Control (OADC) (06.4)
- The OADC shall determine the relative position of the spacecraft.
- The OADC shall determine motion relative to the Earth.
Bus: Structures/Thermal Protection System (TPS) (06.5)
- The structure shall maintain integrity and support subsystems throughout the mission life.
- The TPS shall maintain satellite and subsystems temperatures within ranges of normal
operation.
Mission Operations (07.0)
- The Mission Operation shall generate a plan which allows for the success of the mission
requirements.
Launch Vehicle (08.0)
- The launch vehicle shall safely deliver BOSS to a 700 km polar orbit.
Ground Support System (GSS) (09.0)
- The GSS shall be able to downlink and uplink data to the satellite at least once per day.
Mission Success Criteria
The success of the BOSS mission depends on the ability for the spacecraft to utilize sensors for
in-situ measurements of debris particles. There are several aspects which are critical to mission
success including ability of BOSS to receive commands, measure impact, and send debris impact
measurement data to ground station. The mission success criteria is outlined below.
1.
2.
3.
4.
5.
6.
7.
8.
9.
COMM is able to receive commands from ground station.
COMM is able to send received commands to C&DH to be sent to Payload.
Payload is able to deploy and lock DRAGONS array.
OADC is able to position satellite in desired direction for debris impact.
DRAGONS are able to collect impact data, including size, speed and direction of impact, for
particles of up to one millimeter in size.
Payload is able to send impact data of size, speed and direction to C&DH.
C&DH is able to send impact data to COMM system.
COMM is able to send impact data to the ground station.
GSS is able to receive impact data from Satellite.
Critical Mission Milestones
The BOSS project schedule contains several milestones which are used to assess project
management. The following list displays the milestones which are of most interest:
Flight Readiness Review (FRR)
Launch
Initiate De-Orbit
Decomissioning Review (DR)
Q4 2015
Q2 2016
Q1 2018
Q2 2018
9
Concept of Operations
The concept of operations provides an outline for the structure of the mission beginning at and
following launch. Figure 1 below provides a graphic representation of the concept of operations,
and each portion therein is explained in the following sections.
Launch
Power Flight
Computer, Torque
Rods, and Reaction
Wheels
Detumbling
Solar Panel
Deployment
Deorbit
Operations
Payload
Deployment
Systems Check
Figure 1- Concept of Operations for the BOSS project
Launch
BOSS’ operations begin at launch, which is currently scheduled for Q2, 2016. SpaceX’s Falcon
9 rocket has been chosen as the launch vehicle since it is proven and cost effective. Cost analysis
and comparisons to similar launch vehicles can be found in the Budget section. There are
numerous Falcon 9 missions scheduled for 2016 which allows for flexibility should a launch be
canceled or delayed. The Falcon 9 is also regularly launched from Vandenberg. Vandenberg
launches provide access to polar orbits (90º inclination) which is the desired inclination since
they intersect with all other LEO, thus providing access to space debris located outside of its own
orbit. This would ultimately increase the number of collisions, thus fortifying the body of data
collected.
An altitude of 700 km was chosen since this is the approximate altitude in which the 2007
Chinese Anti-Satellite Missile Test (865 km) and the 2009 Iridium-Cosmos collision (792 km)
occurred. These two recent events drastically increased the amount of space debris in orbit and it
would be wise to search for debris near these altitudes. More information and analysis regarding
orbit selection can be found in the Satellite Subsystem Design section under Payload.
10
Deployment & Stabilization
When the launch vehicle has reached the desired altitude of 700 km, BOSS will be deployed.
The undeployed solar panel will begin energy generation to power the flight computer, torque
rods, and reaction wheels. The torque rods will be responsible for detumbling and stabilizing the
satellite by interacting with the Earth’s magnetic field to slow down and eventually stop the
satellite from spinning. After stabilization, the reaction wheels will be responsible for controlling
the attitude of the satellite and position it in the correct orientation.
Following stabilization, the solar panel will deploy for more sufficient energy generation. A full
systems check by the on-board computer will follow solar panel deployment. When the
satellite’s systems check is complete, the payload, consisting of six DRAGON sensors securely
stowed via five hold-down mechanisms, will be deployed through the use of a spring-loaded
hinge mechanism. Detailed information justifying the use of six sensors and how deployment
works is discussed in the Satellite Subsystem Design section under Payload.
Operations and Deorbit
The operations phase will follow the deployment of the payload and is planned to last 698 days.
During this time BOSS will collect and transmit data via a patch antenna to multiple NASA
operated ground stations. The satellite will send payload data and system health information to
the NASA ground stations which will relay the information to a central Bravo Aerospace
Industries (BAI) facility. BOSS will also be able to receive commands via uplink from any
authorized ground station. This is necessary to meet FCC requirements, and allows controllers to
override a failing subsystem and to initiate the deorbit phase.
After completion of the life cycle, GSS will signal the de-orbiting module to activate. The deorbit module for BOSS will be NanoSat Terminator Tape from Tethers Unlimited. It is a onetime deployable tape that will extend to at least 100 meters. Gravity gradient forces will align
the tape with the local vertical direction to be below the satellite. The tape will increase the
aerodynamic drag of the system and will generate electromagnetic drag due to passive
interactions with the Earth’s magnetic field and ionospheric plasma. There will be two modules
mounted by four #4-40 screws on the outside of the spacecraft. The GPS data will determine if
the first has deployed. If the first module has malfunctioned, the second module will deploy.
The satellite will descend from its 700km orbit in under 10 years satisfying the 25 year
requirement. Figure 2 shown below provides 4 different curves at inclination angles of 28.5, 50,
75, and 98.5, with BOSS performing closely to the 98.5 curve (inclination of BOSS is 90
degrees). BOSS will continue to gather data autonomously as long as possible or until re-entry,
when the majority of the spacecraft will burn up in the atmosphere. The satellite has not been
deemed a threat upon re-entry.
11
Figure 2- De-orbit time versus altitude for a 180kg spacecraft with a 150m Terminator Tape as a function of
inclination, for a 17cm wide tape. (from Tether Tapes Unlimited)
Satellite Sub-system Design
BOSS is composed of seven subsystems, which have been designed to meet their applicable
system requirements. Each subsystem’s design is discussed in detail in this section. Additional
information (including trade study matrices, detailed drawings, and verification/validation
matrices) can be found in the appendix.
Structures
Design Selection
The mission of the structures subsystem is to develop the support and housing of components as
well as allow for the deployment and attachment of external components including solar arrays,
antennas and sensors. In order to perform this, several design concepts were originally conceived
and eventually those options were whittled down to a final design choice. The strength and
torsional rigidity of the frame, the thermal expansion of the various structural components, the
ability of each design to burn in de-orbit, and the complexity of the designs were measured either
quantitatively or qualitatively. Space heritage, the mass and volume of the structure, and the cost
and manufacturability of each design option were also taken into consideration.
The conclusion was that the best design would be comprised of a 0.6m x 0.5m x 0.6m
rectangular frame made of 6061-T6 aluminum alloy with 4 truss-like side webs on the top,
bottom, and two opposite sides. The frame and webs would each have a thickness of 2 cm.
Honeycombed aluminum side panels with a thickness of 7 mm would surround the frame to
12
protect the internal components of other subsystems from the harsh environments observed in
lower earth orbit. These panels were chosen to reduce the mass observed from pure 6061-T6
aluminum panels while simultaneously maintaining the strength of the design.
The frame is comprised of 12 rectangular rods which have a 2 cm x 2 cm cross-section. These
rods are then welded together (using GTAW welding). The webs are used to support the frame
on the top, bottom, front and back sides of the satellite. We chose to use these supports on the
top face to support the DRAGONS sensors and on the bottom face to support the solar array and
antenna. The webs on the front and back faces of the satellite are in place to aid in attaching
several internal components as well as the IMU/reaction wheel shelf. A total of 32 brackets will
be manufactured from 6061 T6 aluminum along with 256 M6 x 1.0 , 12 mm long screws to
attach the support webs to the frame using threaded holes drilled into the frame. Also, 64 M6 x
1.0, 16 mm long screws will be used to attach the panels to the webs and frame with inserts
being used to go through the panels. Pictures of key structural components and a model of the
interior of the satellite, without its internal components, can be seen in figures 3-5.
Figure 3 – Picture of outer frame of satellite
13
Figure 4 – Picture of web used in satellite structure
Figure 5- Internal view of the frame with visible webs and fasteners present.
Interfaces
A critical component of the structures subsystem involves the numerous interfaces that it has
with other subsystems. The structures subsystem was responsible for housing the internal
14
components of each subsystem as well as creating and maintaining a frame that would support
deployable or exterior components for each subsystem.
Structures will maintain physical interfaces with OADC, Payload, COMMS, EPS, TPS and
CD&H. These can all be seen below in table 1. The majority of interfaces are internal. These
are the IMU, reaction wheels and torque rods used for OADC; the transmitter and receiver for
communications; the lithium battery and PCM/BCM for EPS; 7 internal thermistors (thermal
sensors) for TPS and the software used to store data for C&DH. The external interfaces involve
the deployment of the DRAGON sensors for payload, sunsensors and the de-orbit module for
OADC, the solar array for EPS and a thermal coating layer (MLI) for TPS.
Table 1. The different structural interfaces with other subgroups
OADC
Number of
Interfaces
6
Interface (s)
Payload
7
DRAGONS sensor deployment mechanism, control/data box supports
(one for each individual sensor)
Comms
3
EPS
3
Support/attachment for antenna, attachment for the
receiver/transmitter
Solar Panel deployment mechanism/attachment to frame, lithium
battery, PCM/BCM
TPS
C&DH
8
1
Housing and support for de-orbit modules, gyroscope, reaction
wheel, torque rod and external sun sensors
MLI coating and internal thermal sensors
Computer board housing
Load Analysis
In performing the loads analysis on the structure of the satellite, a few assumptions were made.
First, it was assumed that the analysis was to be performed at launch conditions to show that the
satellite could survive launch. Thus, the loads were all assumed to be at 16g. It was also assumed
that the bottom of the satellite was to be attached to the fairing of the launch vehicle. Using this
fact, it was assumed that the bottom of the satellite was fixed during the analysis. The final
assumption is that, while folded during launch, the DRAGONS sensors would have two-thirds of
their mass supported by the top panel and web and one-sixth of their mass supported by each of
the side panels as the sensors latch onto the sides while folded.
In order to perform the analysis, the structure of the satellite was re-created in a single
SolidWorks part file to take advantage of its simulation software. In doing this, the key elements
of the design were retained but a few finer details (such as some of the screw holes) were
suppressed to allow the software to run. Thus the results obtained are not exact but are a good
approximation.
Then the loads acting on the various parts of the structure of the satellite had to be defined. These
were taken to be the loads from the internal and exterior components of the satellite that were
15
supported by the structure (the DRAGONS sensors, solar arrays, reaction wheels, etc.). Table 2
below shows the load breakdown on each of the structural parts.
Table 2. Mass/load breakdown for structural load analysis
Part
Supported Components
mcomponent
Fload  16mcomponent g
Top Panel
Front Panel
Bottom Panel
Back Panel
Right Panel
Left Panel
DRAGON sensors
None
Solar Array, Antenna
De-orbit Modules (2)
Sun sensor, DRAGON sensors
Sun sensor, DRAGON sensors,
Torque Rods (2)
8.46 kg
0 kg
3.504 kg
3.2 kg
2.397 kg
4.049 kg
1327.9 N
0N
550 N
502.3 N
376.23 N
635.53 N
Top Web
Front Web
DRAGONS electronics boxes
Transmitter, Receiver, Torque
Rod, IMU
C&DH computer, EPS Battery
and components, Receiver
Reaction Wheels
Reaction Wheels
IMU
1.411 kg
4.051 kg
221.47 kg
636 N
4.07 kg
639 N
2.525 kg
7.575 kg
9 kg
397 N
1189 N
1413 N
Bottom Web
Back Web
Bottom Shelf
Top Shelf
The results of the analysis (performed through SolidWorks) are presented in Figures 6 below.
Figure 6. Results of stress analysis on outer panels of satellite
16
Figure 7. Results of stress analysis on shelving in satellite
Figure 8. Results of stress analysis on top and front webs of satellite
17
Figure 9. Results of stress analysis on back and bottom webs of satellite
The results of the study show that the maximum stress felt in the structure under these loading
conditions is 143.892 MPa and that the maximum displacement of any part of the structure
during launch will be 5.78 mm. The study also states that both of these maximums will occur in
the interior shelving of the satellite which is understandable since the shelving was handling the
largest loads during the simulation.
In analyzing the results, one key point must be made. While the above pictures show that
SolidWorks believes the yield strength of 6061-T6 aluminum alloy to be  y  55 MPa , [20] and
many other reliable sources state that 6061-T6 aluminum alloy has two values of yield strength,
namely  y  240 MPa and  y  140 MPa (the tensile and shear strengths respectively). Looking
at the above results, we see that the shear forces acting on the satellite all appear to act on the
front, right, left and back panels so as long as none of these stresses reach the maximum value,
we are safe. It can be seen in Figure 4 that none of the panels even leave the blue range (thus
their stress is less than approximately 25 MPa) so the structure can be assumed to survive launch
with a minimum factor of safety given by:
FOS min 
y
240 MPa

 1.67
 max 143.892 MPa
Under this assumption, the structure will be able to support the satellite once in orbit since the
loads experienced there are significantly less than launch loads (the loads will be in fractions of a
g-force instead of 16g).
18
TRL
The structure of the satellite is fabricated out of 6061-T6 aluminum alloy with an external
coating of MLI (multi-layer insulation). This alloy and coating has been used in space
extensively before and thus is TRL 7+.
In regards to the GTAW welding being performed to fabricate the frame, there are NASA
practices that specify the process of using GTAW on spacecraft. Specifically, NASA practice
No. PD-ED-1205 [21] certified GTAW for use on the space shuttle among other missions (a link
to the sheet can be found in the references section). Thus it is assumed that these welds have
been flown into space previously and are TRL 7+.
Risk Analysis
Risk analysis for the structures sub-group is crucial to the survival of the satellite. If the structure
of the satellite or any of the components that are holding other components in place fails during
the mission, there is an overwhelming chance that the entire mission will fail with it. The major
risks that structures has to deal with are failure of the structure during launch, debris impact on
the satellite puncturing the panels or fatigue from oscillating thermal stresses causing the
structure to fail during the mission life.
In order to minimize the chances of these risks occurring, multiple steps have been taken. First,
the loads analysis performed by the team assumed that all loads acting on the satellite were
launch loads. Thus, it was shown that all of the structural components of the satellite can
withstand the launch process. To further minimize the chance of a launch failure, the structure
will undergo vibrations testing after it is fabricated to further validate its ability to survive
launch.
To account for the potential of debris rupturing the panels, we have designed the satellite to have
minimal loading on its panels once the DRAGONS sensors are deployed either through using
webs as additional support or purely through a lack of attached components to the panels (for the
side panels). This allows the panels to withstand higher impacts by allowing them to absorb more
energy. Finally, in order to minimize the possibility of fatigue causing the structure to fail, the
satellite structure will undergo thermal fatigue testing after fabrication to ensure its survival
throughout the mission.
Payload
STK Models for Estimated Impacts
The purpose of the mission is to provide in situ measurements of micrometeroids and orbital
debris (MMOD) in Low Earth Orbit. Systems Tool Kit (STK) software was used to obtain a
preliminary estimate of the amount of impacts to be expected for a given sensor area. However,
successful completion of the mission will enable more accurate models to be created because
actual in situ measurements will be taken.
From earlier trade studies, it was determined that six 0.5 m by 0.5 m DRAGONS units, totaling a
sensor area of 1.5 m2 would be used to detect the debris and take measurements. It was also
19
determined that the satellite would operate at a 700 km orbit at a 90 degree inclination for
optimizing the number of impacts. A model was created for the two year mission for comparison
to other possible cases of orbit size and inclination for a given sensor area. The data for four orbit
sizes (700 km, 750 km, 800 km, and 850 km) and three inclinations (0°, 45°, and 90°) are
provided below, along with a plot showing the trend of the results.
Inclination (Degrees)
0
0
0
0
45
45
45
45
90
90
90
90
Orbit Size (km)
Sensor Area (m2)
Number of Impacts
700
1.5
237
750
1.5
248
800
1.5
254
850
1.5
258
700
1.5
261
750
1.5
273
800
1.5
280
850
1.5
284
700
1.5
361
750
1.5
377
800
1.5
387
850
1.5
393
Figure 10. Data collected for the number of impacts at various orbits and inclinations.
Figure 11. Plot of the estimated number of impacts versus orbit size for various inclinations.
20
As can be seen from the data above, it is predicted that as the degree of inclination increased, the
number of impacts increased, with the 90° inclination having more impacts by a much larger
margin than the other two inclinations. The number of estimated impacts also increased as orbit
size increased, but only by a small amount. For optimizing the number of hits, it can be
concluded that the 90° inclination was an appropriate selection. The number of impacts relating
to the altitudes between 700 and 800 km did not vary enough to say that increasing the orbit by
just 100 km would provide a significant increase in the estimated number of impacts. Thus, the
700 km orbit selected will still provide sufficient data for the mission.
Overall, these models provide an estimate as to what can be expected in LEO for the purposes of
our mission, but it is understood that this area of study still has many unknowns. A mission of
this type has never been conducted before, so the data collected during the mission will
significantly improve existing models like the ones obtained from STK. Therefore, the mission
will benefit users who need more accurate predictions of small MMOD and can be a factor in
improving other missions in LEO in the years to come.
Final Design
The mission of the Bravo Orbital Space Satellite (BOSS) is to conduct in situ measurements of
micrometeoroids and orbital debris that cannot be detected from ground based systems. The
finalized design of the payload incorporates components that optimize the number of MMOD
impacts while remaining within the constraints of mass, budget, and occupation of a minimum
volume in the launch vehicle fairing. Figure 12 displays a schematic of the payload in the fully
deployed position. While the payload is configured in this position, the BOSS will have the
ability to collect data on the number and characteristics of impacts.
Figure 12. BOSS Payload in fully deployed position.
21
Table 3. Bill of Materials of all components incorporated into the BOSS payload.
Item #
p01
p02
p03
p04
p05
p06
p07
p08
Name
DRAGONS Shroud
DRAGONS Control Box
DRAGONS Sensor
Honeybee Robotics
Deployment Hinge
Hold Down Release
Mechanism
Grade 8 Alloy Steel Hex Head
Cap Screw Bolt
Lightweight Aluminum Flat
Washer
Grade F Nylon-Insert Hex
Flange Locknut
Quantity
6
6
6
12
5
130
130
130
An overview of the components used in the payload includes the Debris Resistive Acoustic Grid
Orbital Navy Sensor (DRAGONS) units, deployment hinges, hold down release mechanisms,
and the hardware used for integration to the satellite. A description of each component is as
follows, including any modifications to the components listed.
A single DRAGONS unit consists of a sensor, a control box, and a shroud/frame. Six
DRAGONS units were used on the payload of BOSS, so in total there were six sensors, six
control boxes, and six shrouds that were integrated together to yield the configuration in Figure
12. The DRAGONS units were donated by NASA for the purposes of this mission and will be
provided to Bravo Aerospace Industries as in Figure 13.
Figure 13. Solid model of the unmodified DRAGONS unit frame as provided by NASA to BAI.
22
The frame in Figure 13 that will be provided by NASA will be modified to meet the needs of the
BOSS payload. The NASA DRAGONS unit frame includes four mounting feet and vertical
corner angle members. These mounting feet and vertical corner angle members will be removed
because they will not be necessary to the mounting of the payload to the satellite, and will not
compromise the structural integrity of the unit. The BOSS modified DRAGONS unit can be seen
in Figure 14.
Figure 14. Modified DRAGONS unit to meet the needs of the BOSS. Mounting feet and vertical corner
angle members were removed.
The purpose of the shroud is to act as a frame for the sensors while also preventing MMOD
particles from impacting the sensor at shallow angles. Impacts at these shallow angles have the
potential to remove an excess of sensor detection area if not prevented.
The electronics box for each DRAGONS unit will be placed inside the main satellite structure
and wired to the various DRAGONS units. The control box measuring ~16.51 cm × 21.59 cm ×
8.89 cm is shown in Figure 15.
Figure 15. Mock up version of control box. Actual control box will be provided by NASA
23
The 50 cm x 50 cm sensor of each DRAGONS unit provided by NASA consists of a resistive
grid, a dual-layer Kapton film, and polyvinylidene fluoride (PVDF) acoustic impact sensors.
Upon particle impact, the resistive measurements on the grid change and the size of the impact
area can be measured from these resistance changes. The dual-layer Kapton film is used to
measure the speed of the impacting particles by measuring the time difference between striking
the two layers. The PVDF acoustic impact sensors provides data on the particle impact location.
By placing the acoustic impact sensors at various locations along the sensor, the decomposition
of signals received will provide data on the location. Each 50 cm x 50 cm sensor consists of four
25 cm x 25 cm quadrants of resistive grids coated with a dual-layer thin Kapton film measuring
25 µm thick. The specifications for each quadrant resistive grid include 1960 parallel lines that
are 250 mm long, 62.5 µm wide, and spaced 62.5 µm apart. 3,11
The twelve Honeybee Robotics Deployment Hinges will be the main components used for
payload deployment. These deployment hinges were selected because of their high stiffness, high
strength, and low part count. It was also necessary for the BOSS to utilize torsion spring-driven
damped deployment hinges to prevent an excess of velocity during unlatch that has the potential
to spin the satellite out of control or make excess use of an attitude control system. The hinges
are equipped with micro-switches that will tell BAI when the payload is fully deployed, and have
a full 180° range of motion. The approximate survival temperature for the hinges is -150 °C to
+150 °C, which is sufficient for the mission. The Honeybee Robotics Deployment Hinges are
flight proven and have been used in the USAF STPSat-1 that was launched in 2007.13 These
hinges are also contracted to be used on FORMOSAT-5 in 2014. Figure 16 displays a picture of
the deployment hinges that will be utilized for the BOSS payload.
Figure 16. Honeybee Robotics Deployment Hinges.[11]
Due to the nature of this project, the CAD model of the BOSS payload did not utilize the exact
model of the deployment hinges from the supplier. A similar model was used for the hinge
components that show the overall mechanism of the hinges. Detailed features including the exact
CAD model were not attainable from the supplier.
24
Five hold down release mechanisms (HDRMs) were also utilized for the payload deployment
process. HDRMs, displayed in Figure 17, maintain the load of the DRAGONS units and hold the
deployable sensors in place. They act as the interface between the main satellite structure and the
DRAGONS units and once activated, allow movement of the hinges into the fully opened
deployment position. These mechanisms are flight proven and will be supplied by Sierra Nevada
Corporation Space Systems (SNC Space Systems). A picture from the supplier is seen in Figure
18.
Figure 17. Hold down release mechanism.
Figure 18. Hold down release mechanism.22
The hardware that will be used to secure the sensor shrouds to the hinges and the hinges to the
satellite body were specifically selected to withstand the space and launch environments. The
range of temperatures that the hardware will experience is -100 °C to 150 °C. All fastener
hardware is rated for -150 °C to 250 °C. The bolts selected are made of grade 8 alloy steel which
25
has a 150 ksi minimum tensile strength. This rating exceeds the loads that these bolts will
experience minimizing any risk of fastening failure. The nuts chosen are grade-F nylon insert
locknuts. They were selected for their properties to be able to withstand heavy vibration during
launch, without the nuts backing out. All fasteners are COTS parts from McMaster-Carr. The
exact sizes needed to fasten the hinges are yet to be released by Honeybee Robotics, but nominal
sizes are expected and easily obtainable in the selected material types.
DRAGONS Units Deployment
Stage 1:
During launch and while the satellite is in the launch vehicle fairing, the satellite will be
configured in a compact, cube-like orientation, as in Figure 19. The compact payload design
minimizes fairing volume, allowing the satellite to ride as a secondary, or even tertiary payload.
A minimum volume reduces the high costs associated with the launch operations for both BOSS
and the fairing payload as a whole.
Figure 19. Stage 1 of the Payload deployment. This figure shows the pre-deployment, most compact configuration
of the DRAGONS units around the main satellite structure.
The satellite is equipped with six 50 cm x 50 cm DRAGONS units, twelve Honeybee Robotics
deployment hinges, and five SNC Space Systems hold down release mechanisms (HDRMs).
During stage 1 of deployment, the HDRMs maintain the load of the DRAGONS units and hold
the deployable sensors in place. Figure 18 displays a schematic of the hold down release
mechanisms. These HDRMs are essentially the interface between the main satellite structure and
the payload that allow for full deployment with minimal additional mechanical components.
26
Stage 2:
During stage 2 of deployment, the first set of HDRMs release the DRAGONS units. The
HDRMs are electrically activated, and once activated release a latch that allows the hinges to
operate. Figure 20 provides an image of the deployment right after HDRM release. The hinges
that are being used for deployment are torsion spring-driven damped hinges and provide the
rotational force to deploy the payload. These Honeybee Robotics hinges have high stiffness and
strength along with a low part count to minimize associated mechanical risks. The hinges are
damped to reduce the velocity after release and prevent additional rotational force that could spin
the satellite out of control or require maneuvers from an attitude control system.
Figure 20. Stage 2 of deployment. This figure displays the payload just after HDRM release.
Stage 3:
Stage 3 of deployment is an intermediate phase in which the first set of three DRAGONS units
are fully deployed. The next set of three DRAGONS units are released by the HDRMs and
continue deploying during this stage. Figure 21 shows the first three DRAGONS units in the
locked position, while the second set of sensors has not yet deployed.
27
Figure 21. Stage 3 of Deployment. Half of the DRAGONS units are fully deployed, while second set are about to
deploy.
Upon continued deployment, the sensors will be in a similar configuration to Figure 22 below, in
which the second sets of hinges are partially deployed.
Figure 22. Stage 3 of Deployment. The second set of DRAGONS units has just begun to open up.
28
Stage 4:
Stage 4 is the final stage of payload deployment. In this stage, the DRAGONS units lock into
place and deployment is complete. In order to be sure of full deployment from the ground, each
hinge is equipped with a micro switch that sends a signal to allow Bravo Aerospace Industries to
know when deployment is complete. The fully open position of the DRAGONS units payload is
displayed in Figure 23.
Figure 23. Stage 4 of Deployment. The payload is in the fully opened position
TRL Advancement Plan
The goal of this mission is to introduce a new technology—the Debris Resistive Acoustic Grid
Orbital Navy Sensor into LEO that will characterize and provide data on small orbital debris.
The mission is being conducted for the purpose of advancing the DRAGONS provided by NASA
to a technology readiness level of 7 or higher. Bravo Aerospace Industries is depending on the
supplier to test the DRAGONS units in relevant environmental settings to advance the product to
a TRL 6. After successful completion of this mission, the DRAGONS units will be flight proven
and will provide data on small debris in LEO that ground based systems cannot detect therefore
advancing the sensor units to the appropriate TRL.
All components listed in the bill of materials for the BOSS payload are commercial off the shelf
(COTS) components. All components provided by the suppliers are flight proven with heritage
and will be tested by the suppliers prior to delivery. The suppliers will provide this test data to
BAI upon delivery of the item, ensuring that the components are at the appropriate technology
readiness level.
29
Thermal Protection
Aluminum Radiation Shield
Subsystem components such as those from C&DH and EPS have delicate hardware whose
performance can be degraded by cosmic and solar radiation. Specifically the PDM and BCM
components of EPS must be protected from exposures greater than 5 kRad while the lithium-ion
battery must remain under 500kRad for its lifetime. To meet these requirements, 7mm thick
honeycomb aluminum panels and a radiation shield container will house these components.
Figure 24. Radiation Shielding Design This shows how the total life time radiation dosage inside a spacecraft
changes as the amount of aluminum shielding increases. 23
The MIR space station was maintained at a near earth orbit at an altitude of 296-421km while
ISO had a highly elliptical geocentric orbit with apogee at 70600 km and perigee at 1000 km.
The curves for radiation dosage relative to aluminum shielding thickness for both of these
satellites are shown in figure 24. A curve similar in shape and between the curves of MIR and
ISO would characterize our satellite at 700km altitude polar orbit. Yearly our satellite
components should be kept to an exposure level of about 1.5kRad for its 3 year lifetime.
𝑇𝑜𝑡𝑎𝑙𝐸𝑥𝑝𝑜𝑠𝑢𝑟𝑒(𝑘𝑅𝑎𝑑)
= 𝐸𝑥𝑝𝑜𝑠𝑢𝑟𝑒 𝑝𝑒𝑟 𝑦𝑒𝑎𝑟 (𝑘𝑅𝑎𝑑)
𝑌𝑒𝑎𝑟𝑠 𝑜𝑓 𝐿𝑖𝑓𝑒
30
To maintain the satellite under 1.5kRad exposure the data presented in figure 24 shows that 6
mm of 6061-T6 aluminum alloy will suffice. The honeycomb panels of the exterior consists of a
honeycomb pattern sandwiched between two layers of 1 mm thick aluminum alloy sheets.
Between the exterior panels of the satellite and the radiation protection container there is also
aluminum webbing and other components, which may absorb radiation. However, the coverage
is not extensive throughout the structure to adequately provide protection. Therefore a box with 6
mm thick aluminum of welded sheet metal construction with a close fit lid will be used. The total
mass of this container will be approximately 5 kg.
Multi-Layer Insulation
To test how well the Multi-Layer Insulation (MLI) protects the satellite from the harsh space
environment, a number of heat transfer equations must be utilized. The space environment
temperature varies depending on where the satellite is and the time of day so it is important that
the MLI provides protection for a wide range of temperatures.
To complete this task, the surface temperature of the satellite must be predicted using:
𝑞𝑎𝑏𝑠𝑜𝑟𝑏𝑒𝑑 = 𝑞𝑒𝑚𝑖𝑡𝑡𝑒𝑑
𝐺𝑠 𝛼𝑠 = 𝜀𝐼𝑅 𝜎𝑇 4
1
𝐺𝑠 𝛼𝑠 4
𝑇=(
)
𝜀𝐼𝑅 𝜎
(1)
where,
𝐺𝑠 = Solar Flux
𝛼𝑠 = Absorptivity
𝜀𝐼𝑅 = emissivity
𝜎= Stefan-Boltzmann constant
Once the outer surface temperature of the MLI is calculated, we can treat the MLI as a plane wall
with each layer having its own thermo and physical properties. Because the MLI is simplified to
a plane wall, the heat transfer rate through each layer is equal throughout the insulation. To put
it simply:
𝑞𝑙𝑎𝑦𝑒𝑟𝑖 = 𝑞𝑙𝑎𝑦𝑒𝑟(𝑖+1) = 𝑞𝑙𝑎𝑦𝑒𝑟(𝑖+2) = 𝑞𝑙𝑎𝑦𝑒𝑟(𝑖+𝑛)
Each heat transfer rate equals
𝑞 = 𝛼𝐺 − 𝜀𝜎(𝑇𝑛+1 4 − 𝑇𝑛 4 ) −
𝑘(𝑇𝑛+1 − 𝑇𝑛 )
𝐿
(2)
Where 𝑇𝑛 equals the outer temperature of the next layer (it also represents the inner temperature
of the previous layer). 𝑇𝑛+1 is the inner temperature of the layer (or the outer temperature of the
next layer).
One can see that MLI is very versatile. The number of layers and their properties all affect how
the MLI will perform. Figure 25 shows the typical makeup of MLI and Figure 26 shows the
absorptivity and emissivity of common MLI materials.
31
Figure 25. Common Multilayer Insulation makeup
Figure 26. Absorptivity and Emissivity of common MLI materials
For the BOSS satellite design, Figure 25 was used to create a preliminary layout for the MLI and
the properties from Figure 26 were used to finalize the design of the MLI. Following the format
of Figure 25, the BOSS satellite’s MLI will consist of:





Outer layer and inner layer of Kapton Film
Beta Cloth to act as a light block
Aluminized Teflon Optical Solar Reflectors which have high emissivity to block sunlight
Netting Spacers and Mylar sheets to reduce conductivity (13 layers in total)
All held together with Elastisol space adhesive and tape
32

The whole MLI will be attached to the structure by hook and pile connectors
Here is an example photo of MLI that is very similar which will be used on the BOSS satellite
Figure 27. Typical MLI makeup.
The golden film on the outside represents the Kapton film backed by aluminum. The solid
metallic layers represent the reflectors. The netlike materials represent the spacers, and inbetween these layers are low conductivity Mylar sheets. To confirm that this layout works, the
MLI was tested when the satellite resided in the bright side of the Earth and on the dark side of
the Earth.
Bright side of the Earth
Using the given values:
𝐺𝑠 = 1371 W/m2
𝛼𝑠 = 0.41
𝜀𝐼𝑅 = 0.75
𝜎 = 5.67 x 10-8
When the satellite is on the bright side of the Earth, the surface temperature of the satellite, using
Equation 1, is:
1
𝐺𝑠 𝛼𝑠 4
𝑇=(
) = 339 𝐾 = 66°𝐶
𝜀𝐼𝑅 𝜎
Using this value and plugging into Equation 2 yields
𝑞 = 𝛼𝐺 − 𝜀𝜎(𝑇𝑛+1 4 − 𝑇𝑛 4 ) −
𝑘(𝑇𝑛+1 − 𝑇𝑛 )
𝐿
This value will be used to find the changes in temperature when heat travels through the
insulation. Substituting the given properties for each layer, the final temperature or the inside
temperature of the satellite will 16°C.
33
Dark side of the Earth
The final temperature of the satellite is calculated the same as the bright side, but the only
difference is the solar flux value: 𝐺𝑠 = 340 W/m2. This changes the initial surface temperature of
the satellite.
1
𝐺𝑠 𝛼𝑠 4
𝑇=(
) = 239𝐾 = −34°𝐶
𝜀𝐼𝑅 𝜎
Following the same procedure as above, using the same given thermal properties, the final
temperature of satellite will be -16°C. With the internal heat generated by the components and
the aluminum box for radiation protection, this is warm enough for the safe operation of the
internal components.
Internal Satellite Thermal Analysis
Many of the heat generating components within the satellite are also contained within radiation
protection boxes and where therefore analyzed as boxes. The thermal analysis was run at the
calculated interior ambient temperatures the MLI leads to of -16ºC to +16ºC. Components with
sensitive temperature ranges where considered to analyze the impact of surrounding components.
Radiation was the main mode of heat transfer with most components consisting of Aluminum
alloy with emissivity of 0.25 and polymers with emissivity of 0.85.
Figure 28. Thermal Analysis at +16 ambient.
34
Figure 29. Thermal Analysis at -16 ambient.
Thermistors
To monitor the thermal health of the satellite seven thermistors will be employed, one mounted
on the interior of each honeycomb aluminum alloy panel and one mounted on the lithium-ion
battery radiation protection box. This placement presents the opportunity to manage components
that are out of their operating thermal range by changing the individual components power mode.
The thermistors chosen have a temperature range -60ºC to +160ºC with tolerance from 0 to
+70ºC : ±0.35ºC. Other useful qualities:
•Excellent Stability
•Flight Heritage
•Robust Construction
•Non ITAR Restrictive
•Flat Surface for Ease of Mounting
•Approved by Prime Contractors
•Flat Aluminum Housing
These thermistors will be mounted onto the satellite using space grade tapes and adhesives. They
are constructed of Matched Glass Coated NTC Thermistor Beads mounted on aluminum housing
and encapsulated in Stycast 2850FT epoxy.
35
Figure 30. Thermistor dimensions.
Figure 31. Thermistor resistance table
Space Grade Wiring
The space environment also impacts component wiring especially when exposed on the exterior
of the satellite such as for the sensors payload. To counter the impact of radiation and thermal
effects the conductor is silver plated copper and the insulation is flexible PTFE that has excellent
outgassing behavior, reinforced mechanical resistance and a temperature range of -100ºC to
+200ºC.
36
Safety and Mission Assurance
There are several considerations for the safety and mission assurance of the Thermal Protective
System. The following outlines how BAI will assure thermal protection remains functional
throughout the mission.
1. Environmental Effects
a. Throughout its time in the atmosphere, the satellite will experience fluctuations in
temperature.
b. Prior to launch, the materials shall be tested at the extreme temperatures of space to
assure that the subsystems will remain at operational temperatures.
2. Degradation and Deformation
a. Structural materials will be tested prior to launch at various temperatures and in
vacuum to verify that the materials will withstand and perform at intended.
3. Contact with debris
a. Multi-Layer Insulation (MLI) would degrade slightly per puncture made but will
continue to function at a lesser extent
b. A high energy impact can sever wires in electrical equipment for active methods such
as heaters.
4. Heat Transfer
a. Subsystems may receive too much heat or not enough, depending on whether or not
the satellite is in the shadow of the earth or exposed to the sun.
b. Radiation box protects vulnerable components like the battery
5. Maintenance
a. Because the lifespan of the satellite is roughly three years, no further maintenance
will be needed.
Risks Management
For the TPS, potential risks include breaks in the MLI, oxidation of the Aluminum Optical Solar
Reflectors, outgassing. This would expose the internal components to the harsh space
environment. Many of the subsystems within the satellite have to be within a temperature range
to survive so to prevent any conflicts, strong insulation and material were chosen. Other risks lie
in manufacturing error, testing error, and thermistor error which can be prevented before launch.
The risk management chart in the “Risk” section details all of the risks and their mitigation
strategies.
Electrical Power System
In order to give the mission a chance at success, the electrical power subsystem must provide
adequate power to each component for the life of the satellite. The electrical power subsystem
needs to generate power from the environment, condition and store that power, and distribute the
proper amount to each component throughout the life of the BOSS two year mission. A total of 6
subsystems have power needs for their components that the electrical power subsystem must
provide. The final design is one that meets all of the above success criteria while helping to stay
within the mass and cost constraints. The final design is an electrical power system that draws
energy from solar arrays, stores energy in a secondary battery, and uses a power conditioning a
distribution unit to deliver power.
37
Design Outline
The electrical power system is designed to collect energy from the sun with a Gallium Arsenide
solar array. This solar array during launch is folded on the bottom side of the BOSS structure.
The array is folded so that half of the total area is exposed, so it acts essentially as a body
mounted panel to provide power during detumbling. After detumbling the solar array will deploy
with a mechanically powered hinge and begin to provide power. The array will always be
directed toward the sun using a biaxial gimbal with positioning commands coming from the
command and data handling flight computer. The BOSS internal electronic components are
made up of a Clyde space Lithium Ion battery and a Surrey made power unit. The Surrey power
conditioning and distribution unit (PCDU) is a battery conditioning module and power
distribution module in one box. The battery conditioning module regulates the incoming power
from the solar array while also limiting the charge and discharge rates of the battery. The power
distribution side of the unit provides the regulated power when it is needed for each component.
The lithium ion battery provides the power needed for hardware components when the BOSS
satellite is eclipsed by the earth.
Figure 32. Above is the hardware schematic for the electrical power subsystem components. Note that everything
within the PCDU is one unit within the BOSS structure.
Power Collection
The BOSS has a single deployable photovoltaic solar array on the bottom side of the structure.
The array is made up of two a square panels 450 mm by 450 mm which fold out using a spring
powered hinge. The 405 mm2 of Gallium Arsenide solar panels provide 79.5 watts of power at
the end of their life. After 76.5 minutes of Sun exposure, the array captures 101.3 watt∙hours of
energy. This is enough to provide the 82.9 Watt∙hours per orbit needed to charge the Lithium Ion
38
battery and provide power to electrically powered components. The solar array purchased from
Clyde Space has a total efficiency of 28.3% and only has a total mass of 1.84 kg.15 The power
calculation from the solar array includes all loses from cell degradation, angle to the sun loses,
and inherent degradation.
Figure 33. Clyde Space Solar Array with Gimbal Assembly15
In order to assure power to all subsystems, the solar array will be assembled with a biaxial
gimbal purchased from MOOG.14 The command and data handling flight computer will provide
input commands to the gimbal based on sun sensor readings from OADC. The result is a solar
array that tracks the sun for the most power input. The gimbal itself requires 10 Watts of power
to move when the array is collecting and adds an extra 1.2 kg of mass.14 The gimbaled and
deployed solar array offers a better power solution than a larger area of body mounted solar
arrays because of the need for the satellite core area for mounting the Dragons payload and
OADC equipment.
Figure 34. The Type 11 Biaxial Gimbal used to control the solar array on the BOSS.14
39
Energy Storage
The battery is designed to provide power to components that need to run during the periods when
the BOSS is eclipsed by the earth and no power can be drawn from the solar array. A lithium ion
secondary battery was selected from Clyde space because of their experience with powering
CubeSats. A lithium ion battery is the obvious choice with a strong space heritage, a high energy
density, and the BOSS limitations on mass. The total energy need in the 21.5 minute dark period
is 13.2 Watt-hours. The battery must be charged and discharged every orbit. Therefore to reduce
damage to the battery, it must not be overcharged at all or over discharged by more than 30
percent.17 There result is a required battery capacity of 51 Watt-hours, including efficiency
losses. A Clyde Space 90 Watt-hour CubeSat battery meets this requirement at a mass of 0.768
kg.17 By over sizing the battery, we assure a power positive system, meaning the BOSS will
always have more power available than it needs.
Figure 35. The 90 Watt∙hour Clyde Space Lithium Ion Battery used on the BOSS17
Power Conditioning and Distribution
The battery and solar array give a method of collecting and storing power, but the EPS requires
components to regulate and distribute this power. The Surrey made Power System 150 PCDU
provides the medium between collecting and distributing. The Power System 150 is comprised of
a battery conditioning module (BCM) and power distribution module (PDM). The BCM is the
interface between the solar arrays and battery.16 This power system provides six battery charge
regulators (BCRs) to regulate the amount of power charging the battery from the solar array. The
BCRs covert the input power to the nominal 28 volt bus needed to charge the battery and
interface the PDM.16 The BCM also has a battery charge monitor which protects against
overcharging and discharging by supplying signals to the BCRs to turn on and off input or to the
OV shunt.18
The power is distributed to the BOSS components via the PDM of the Power System 150 unit.
The PDM takes commands from the command and data handling flight computer to determine if
each of the 28, 28 volt switches and 16, 5 volt switches are on or off. The PDM also provides
over current protection for each output switch.
40
Figure 36. The Surrey Power Unit 150 PCDU used on the BOSS.16
Power Budget
It is critical to mission success that the power budget is accurate. The budget determines if the
solar arrays are large enough to both provide power to components and charge the battery while
the array is in the sun. The budget also includes how much energy each subsystem uses during
the eclipse period to determine the size of the Lithium Ion battery. Tables 4 and 5 illustrate the
overall power budget, while the third shows the power needs on the dark 22 minute period.
Table 4. Power budget for BOSS.
Component
Command and Data
Handling CPU
Dragons Sensors
GPS
IMU
Sun Sensors
Torque Rod
Reaction Wheel (x3)
Solar Array Gimbal
Communications
Thermocouples
Battery Charge (with
losses)
Total Need
Normal Power
Mode (W)
5 Constant
Secondary Power
Mode
9 for Downlink
Energy Need
(W∙hr)
8.61
6 Constant
1.6 Constant
5 Constant
0.2 Constant
2 As Needed
4.5 As Needed
1 Per Hit
None
None
None
None
Mode 1: 8.4, Mode
3: 39 as Needed
None
15 Downlink
None
9.80
2.61
8.17
0.33
3.27
27.6
10 as needed
1.5 Uplink
2 Constant
1.27
4.12
3.27
13.89
82.89
The system is considered power positive because the solar array generates 101.3 watt∙hours of
energy which exceeds the needed 82.9. The EPS will always generate more power than is
consumed. Notice that OADC demands the most power for an active attitude and orbit control.
41
Table 5. Component energy needs.
Component
Battery Capacity Need (W∙hr)
1.793
Command and Data Handling CPU
2.153
Dragons Sensors
0.574
GPS
1.793
IMU
0.072
Sun Sensor
0.717
Torque Rod
4.842
Reaction Wheel
0.5378
Communications Uplink
Total
13.199
The number of components running in the eclipse period is kept to a minimum to reduce the
mass and cost of the battery. The 90 watt∙hour Clyde Space battery is well above the needed 51
watt∙hours needed to ensure the 30 percent depth of discharge. It also means the battery storage
is power positive. The energy storage is always more than the BOSS will consume.
Figure 37. Above is the wiring diagram from the power distribution module to all the powered components. A
larger version of this figure is in the appendix.
Configuration
The two panel deployable solar array and gimbal is mounted to the bottom side of the BOSS.
The single spring powered hinge is mounted to the gimbal and center of the edge of the bottom
side opposite the deployed Dragons sensors. The PCDU and battery are positioned inside the
structural core of the BOSS for thermal and radiation protection. The spec 80 FlexLine space
grade wiring for all components is provided by TE connectivity. The wiring must be space grade
and tested to avoid out gassing. The wiring runs from each power relay output of the PDM to
each of the components needing power. The total mass of the assembled electrical power system
42
is 8.73 kg. Each component requiring power is wired directly to the PDM of the PCDU. The
wiring diagram below provides all the components connected to the PDM.
Risk Analysis
Electrical power failure is not acceptable and will result in total mission failure. To avoid this, all
failure modes have been considered and designed around. First is failure at the battery. Batteries
inherently degrade over time. To avoid significant losses, a short satellite mission life and
cyclical charge tests eliminate the risk of failure. Over charging and discharging of the battery is
eliminated as a failure mode by designing a large enough battery for eclipse power and including
the charge monitor and over voltage shunt in the BCM. Thermal and structural damage to the
battery is designed around by the thermal and structural subsystems.
Solar array failure is the most likely source of electrical power failure. Deployment and solar
tests by the manufacturer reduce the risk of total failure. Debris could damage part of the array
reducing the total power it can absorb. To reduce the risk associated with debris, the array is
divided into 6 segments that are wired into separate BCRs. This way the loss of one segment of
the array does not cause total failure.
Gimbal testing by the manufacture reduces the risk of loss of mobility of the solar array.
However, this loss of mobility will only result in a reduced power state. The PDM provides over
current protection to all components to eliminate the risk electrical damage. Failure from
electronic connections is eliminated by extensive testing and the use of space grade wiring
harnesses.
Interfaces
The interface between the electrical power components and the command and data handling on
board computer is the most critical interface. Without the commands going to and from these
components, power would not be provided at the correct time and electrical components could be
damaged. The PCDU interfaces with the command and data handling using an RS 422 connector
through which it receives necessary commands. The interface with command and data handling
is summarized in the tables in the appendix. All interfaces with other subsystems are controlled
through the PDM which determines if the component is on or off and the nominal voltage.
Wiring runs from each power relay output on the PDM to each of the components that require
power. These relays are used to switch on and off components when power is needed or not
needed for specific components. Spec 80 FlexLine Wire from TE connectivity is used to route
the power from the RS 422 connectors at the PDM to the components requiring power all of
which are shown in figure 6 above. The spec 80 FlexLine wire is space grade wire insulated with
flexible modified radiation cross-linked ETFE polymer rated from -65 to 200 degrees Celsius.
The wires will be bundled coming out of the PDM and when possible. Also a ground cable will
be run to maintain electrical continuity between structural elements.
Orbit Attitude Determination and Control
Design
An Inertial Measurement Unit (IMU) is a device that senses and calculates the spacecraft’s
orientation and rates by measuring the angular acceleration and angular velocity. The angular
43
acceleration is measured using accelerometers, typically one per each axis for a total of three.
The angular velocity is measured using gyroscopes, again with one for each axis, or a single,
three-axis gyroscope. Normally, IMUs calculate the position of the spacecraft by integrating the
measured angular rates, knowing the time-step. However, such systems will become increasingly
inaccurate over time as error accumulates due to the noise and bias of the sensors, as well as the
integration method.
To remove such error, the IMU proposed is to use a new configuration based on a paper by
Vivek Nagabhushan to reduce the bias and noise of the angular acceleration and angular velocity
measurements.10 In terms of the physical configuration there are six accelerometers, with two on
each axis, to allow better measurement of the angular acceleration by knowing the relative
distance between the two accelerometers.
The measurements are then run through a Kalman filter which will iteratively calculate more
accurate (bias-free) angular rates from the measured values. The Kalman filter process and
equations are described in full by Nagabhushan.10 Not mentioned specifically in the reference
material is the number of iterations to use for each measurement, it has been determined that the
filter will run at a rate proportional to the time step considered in the calculations and the refresh
rate of the sensors. So for a refresh rate of 50 𝐻𝑧 and an internal rate of 5000 𝐻𝑧, the filter will
run 100 iterations for each measurement. There will also be a dedicated 100 𝑀𝐻𝑧 CPU in the
IMU itself to run the filter and send the processed data to the satellite’s main CPU to store.
The specific components of the IMU referenced below can be seen in the (table or appendix).
The accelerometers on the x-axis are positioned in the direction of the y-axis, the y-axis in the
direction of the z-axis, and the z-axis in the direction of the x-axis. The relative distance between
the accelerometers on each axis will be 130 𝑚𝑚 to provide enough distance for the filter to
properly work on removing noise from the angular velocity at least. However, noise will still be
present in the angular acceleration measurements for low-amplitude oscillations, which are
expected. The chosen accelerometers are the QA3000-010 from Honeywell, and provide low
bias and intrinsic noise.
The gyroscope configuration is the same as any commercial IMU, with three ring-laser
gyroscopes positioned along the three orthogonal axes. The gyroscopes may have an internal bias
and may accumulate additional bias over time, the Kalman filtering acts to remove this bias. The
chosen gyroscopes have a bias stability of 0.0035 𝑑𝑒𝑔/ℎ𝑟, which is within acceptable ranges for
the Kalman filtering.
The IMU was simulated using the above components in order to determine the optimal
positioning of the accelerometers. In order to reduce the error in the angular velocity to
acceptable levels, it was determined that the distance between accelerometers should be at least
10 𝑐𝑚, with additional distance reducing the error further. To fit within a certain frame based on
the dimensions of COTS IMUs, the distance between accelerometers was chosen to be 13 𝑐𝑚,
with one placed 2 𝑐𝑚 from the origin of the axes and the other at 15 𝑐𝑚. This configuration
provides a mean error of less than 0.000005 𝑟𝑎𝑑/𝑠 in the filtered angular velocity, which
corresponds to less than 0.1% error. In terms of angular acceleration, the bias is removed in the
filtering, but the noise still prevails. For small angular accelerations as expected during
44
operation, the noise will be significant, however, the general form of the angular acceleration is
still visible and can be determined by taking a mean of the still-noisy measurements after
filtering.
Kalman filtering of the measurements in this configuration provides a clear advantage in
accuracy over other COTS IMUs. The physical development of the structure of the IMU and
further testing are at this stage planned to be performed by University graduate students, with all
components of the IMU being supplied by BAI. This IMU will be a step forward in capabilities
of small satellites. However, should the new configuration not meet requirements it will still be
able to function as any other IMU. Therefore, this new configuration will not pose a risk to the
overall mission, and will only add a new component to the mission as a secondary objective.
In determining the required positional accuracy of the system, the maximum distance of the
DRAGONS payload was considered along with its provided sensitivity of 150 𝜇𝑚. Knowing the
internal dimensions of the satellite and the length of the extended sensors. The length at which
the sensors must be able to discern the location of impact is approximately 1.3 𝑚. Since the
sensitivity was not specified total or in any direction, the given sensitivity was halved to 75 𝜇𝑚
to be safe.
At a distance 𝑥 along a stationary axis from the center of the satellite, the distance 𝑥 ′ on the
sensor at an angle Δ𝜃 from the initial axis is calculated by
𝑥′ =
𝑥
cos(𝛥𝜃)
So, for the sensor to be able to distinguish impacts, it is required that
𝑥 − 75 𝜇𝑚 < 𝑥 ′ < 𝑥 + 75 𝜇𝑚
which can be further simplified to
𝑥
+ 75 𝜇𝑚 > 𝑥
cos(𝛥𝜃)
thus,
𝑥 (1 −
1
) < 75 𝜇𝑚
cos(𝛥𝜃)
Using the length of 1.3 𝑚 for 𝑥, the equation then becomes
1.3 𝑚 (1 −
1
) < 75 𝜇𝑚
cos(𝛥𝜃)
and simplifying,
45
Δ𝜃 < 0.6154°
So, essentially the satellite requires a positional accuracy of less than 0.5°, and by using
Modified Rodrigues Parameters to determine position, this is possible9.
Figure 38.1a. The actual angular acceleration for each axis (x,y,z) in blue, with the filtered angular acceleration in
red. The noise from measurements is still present, but follows the trend.
46
Figure 38.1b. The error in angular acceleration for each axis (x,y,z). The unfiltered angular acceleration is in blue
and the filtered angular acceleration in red.
Figure 38.2a. The actual angular velocity for each axis (x,y,z) in blue, with the filtered angular velocity in red. The
filtered velocity is barely visible above, so the noise and bias have been adequately removed.
47
Figure 38.2b. The error in angular velocity for each axis (x,y,z). The unfiltered angular velocity is in blue and the
filtered angular velocity in red. The error in the filtered angular velocity is much less than the unfiltered.
The other attitude and orbit determination components are two sun sensors and a GPS receiver.
The two sun sensors will be placed on the outside of the satellite and on sides opposite one
another that face outward, perpendicular to the velocity vector. The sun sensors have an accuracy
of 1° and a 90° field of view. The accuracy of the sun sensors will allow the solar panels to track
the sun for the majority of the time from their data alone, and the rest of the time by
extrapolating the position either from the sensors or by the power input. The GPS receiver is able
to get the orbital position and velocity of the satellite with typical accuracies of 10 𝑚 and
0.15 𝑚/𝑠, respectively.
Attitude Control
For attitude control, it is desired for the outward normal of the DRAGONS to remain within 5°
of perpendicular to the velocity vector of the satellite to achieve the maximum number of
impacts on the sensors. It is also desired that the satellite have a nadir pointing accuracy of less
than 5° to accommodate maximum impacts on the sensors and allow the satellite to properly
communicate with ground. The chosen 100SP-O reaction wheels can provide a maximum torque
of 0.11 𝑁 𝑚, which is easily capable of orienting a satellite of this size. Torque provided to the
satellite can be simplified for considering the torque acting along a single axis at the center of
gravity as
𝑇 = 𝐼𝑎
in which 𝐼 is the moment of inertia about an axis. Considering 𝐼 = 0.155 𝑘𝑔 𝑚2 , the maximum
acceleration that can be provided to the satellite along that axis is
𝑎=
0.11 𝑁 𝑚
𝑟𝑎𝑑
= 0.7097 2
2
0.155 𝑘𝑔 𝑚
𝑠
However, operating any reaction wheel at maximum torque is very power-consuming, requiring
13 𝑊. So the goal is to operate at around 5000 𝑟𝑝𝑚 for most control instances, which requires
only 2.8 𝑊, and should provide a torque roughly 20% of maximum, which gives
𝑎𝑜𝑝𝑡𝑖𝑚𝑎𝑙 = 0.1419
𝑚
𝑠2
This optimal acceleration will be sufficient for normal operation to keep the satellite aligned as
desired.
The actual setup will position the reaction wheels as close to the center of gravity as possible,
with four reaction wheels in a tetrahedral configuration to provide three-axis control as well as
redundancy from the fourth wheel. The three torque rods will be placed on three of the inner
edges of the satellite to allow for momentum dumping from the reaction wheels and to avoid
disturbing other systems. The torque provided by a single Andrews TQ-40 torque rod can be
calculated by the equation19
𝑇𝑚 = 𝐷𝐵
48
in which 𝐷 is the dipole of the torque rod, and 𝐵 is the magnetic field of the Earth. For the
chosen torque rods, the nominal dipole is given12 as 42 𝐴 𝑚2 , and the magnetic field of the Earth
in a polar orbit is
𝐵=
2𝑀
𝑅3
in which 𝑀 is the magnetic moment of the Earth and 𝑅 is the radius from the center of the Earth.
For a polar orbit, then
𝐵=
2(7.96 × 1015 𝑇 𝑚3 )
(7071 𝑘𝑚)3
which then gives
𝑇𝑚 = 42
2(7.96 × 1015 𝑇 𝑚3 )
= 0.0019 𝑁 𝑚
(7071 𝑘𝑚)3
This torque is two orders of magnitude less than the maximum torque provided by a reaction
wheel, but only one order of magnitude less than the expected torque needed by any reaction
wheel during normal operation. This will allow for sufficient momentum dumping from the
reaction wheels in a relatively short amount of time compared to other torque rod options.
Risk Analysis
The Honeywell GG1320AN Ring Laser Gyroscopes (RLG’s) in the IMU generate some random
walk and bias stability throughout its life cycle that accumulates as error and failure towards the
end of its lifetime. Random walk at 0.0035deg/hour and bias stability at 0.0035deg/root-hour is
typical for the gyroscope. The design of the IMU incorporates Kalman Filtering (commonly used
in inertial systems) which will filter the excess noise in the system that may appear from these
factors. The mission length is 2 years and the lifetime of the RLG’s is about 7 years so random
walk and bias stability should not become an issue. The RLG’s may be susceptible to jitters
which may skew the data but this can be tracked and should not pose a major issue to data
collection. If one RLG fails, the spacecraft will lose attitude data on an axis. The IMU will
consist of 6 Honeywell QA3000-010 accelerometers with two on each axis used to measure the
difference between them. If one accelerometer fails, the other accelerometer will continue to
take data and operate the same way a normal IMU would work.
Two Surrey 2-Axis sun sensors used in this system to take data. If one sun sensor fails, the
second one will continue to operate with little loss in positional data. Issues may appear with the
data upon a single point failure. The spacecraft could not be facing in the direction of the sun or
there could be reflection off of another satellite or space debris that could skew the data.
If the Surrey SGR-20 GPS fails during the mission, direct position data for the satellite will be
lost. In this case, the path of the satellite can be estimated by established data. This data will
include the spacecraft’s current orbit and drag analysis based on the area affected and the
spacecraft’s altitude. Accelerometer data will provide velocity, which can be integrated to
49
provide the position of the spacecraft over time. The IMU will detect collision of the spacecraft
and track its position change from the impact.
Failure of a Surrey 100 SP-O Microwheel (Reaction Wheel) causes a loss of one axis of control.
The fourth redundant wheel will be used in place of the failed wheel to provide accurate attitude
control for the spacecraft. The redundant wheel ensures that the spacecraft will maintain a
pointing accuracy of 5 degrees.
If one Andrews TQ-40 Dual Wound Torque Rod fails, only the momentum dump process for the
reaction wheels will be affected meaning there will be less stiffness during momentum dump.
The two remaining torque rods will change the attitude of the spacecraft to provide the stiffness
for the momentum dump. Each torque rod is dual would so if a single winding is malfunctioning,
the other winding will be used. Both windings may be used to provide greater stiffness during
each momentum dump. The system may be affected from sign interference by the torque rods
but they will not be used extensively so it should not pose an issue to the operation of the BOSS
spacecraft.
If the Tethers Unlimited Nano-Sat terminator tape should fail, a second module will be deployed.
The GPS will track the altitude of the spacecraft after mission completion to ensure the deorbiting module has deployed and is working properly.
Technology Readiness Level
IMU Testing
The in-house IMU is currently at TRL 3 as Kalman filtering is a proven method of obtaining
highly accurate data from the Honeywell QA3000-010 accelerometers and the Honeywell
GG1320AN gyroscopes. The advancement plan will consist of:
To TRL 4:
 Accelerometers and gyroscopes attached to breadboard or testing circuit board to receive
data
 Setup on turntable for controlled (known) angular acceleration/velocity
 Data sent to computer for processing (Kalman Filtering)
 Data compared to expected accuracies
To TRL 5:
 Setup tested again and restricted to flight computer processing capabilities
 Include small impulses
 Compare with simulations
To TRL 6:
 No longer on breadboard, but fully-functional prototype IMU configuration
 Test again in same conditions as previous
To TRL 7:
 Component either fully integrated into satellite or left separate
 Tested in conditions closer to the actual operational environment
 Parabolic flight tests
 Compare to reported responses from reaction wheels
The IMU will advance to TRL 9 with the flight of the BOSS spacecraft.
50
The Surrey 100 SP-O Small Satellite Microwheels are at TRL 7. All necessary vibration and
thermal testing have been conducted and the product is ready for delivery. The TRL will advance
to TRL 9 with three scheduled flights by early 2014: TDS1 (2013), KAZ MRES (2013), and
DMC3 Constellation (2014).
The GPS Receiver is at TRL 9. It has undergone vibration testing to NASA GEVS acceptance
and undergone thermal cycling. The Surrey SGR-20 Space GPS Receiver has 65 years of total
flight experience with 20 units flown.17
The Surrey 2-Axis sun sensors are at TRL 9. They have undergone vibration testing to NASA
GEVS acceptance and undergone thermal cycling. The sun sensors have 63+ orbit years to
include the FASAT-Bravo, Uosat-12, TopSat, and DMC-4 satellites.16
The three Andrews TQ-40 torque rods are at TRL 9. They have had all necessary testing and
have flight history. They were sold and used between 2007 and 2012 by Sinclair Interplanetary
until transfer to Andrews Space.14 Three units were launched in November of 2010 and 69 units
have been delivered.18
The de-orbit module is at TRL 7. The CubeSat module has undergone space deployment testing
and has been successful. The Tethers Unlimited Nanosat Terminator Tape works under the same
concept. It has been awarded contract for delivery in Q2 of 2014.15 The advancement plan will
be the use and documentation of the NanoSat Terminator Tape for flight which will advance it to
TRL 9.
Interfaces
The OADC subsystem is primarily controlled by Electrical and Mechanical Interfaces, the
electrical interface being more critical of the two. Failure of any single electrical interface could
result in massive failure for the rest of the subsystem and corrupted data in the worst case
scenario. The Electrical interface is run by small interactions between Attitude Control (GPS,
IMU, 2 Sun Sensors), Attitude Determination (4 Reaction Wheels, 3 Torque Rods) and the flight
computer.
The following is a list of all the component interfaces.
1. Electrical
A. Attitude Determination (GPS, IMU, and Sun Sensors) provides position, velocity,
acceleration, angular velocity, angular acceleration, and attitude data to the flight
computer to be sent to the proper parts of the satellite that need them. i.e. position data is
received by Ground Systems.
B. The flight computer provides commands to the Attitude Determination system.
C. The flight computer provides commands Attitude Control (4 Reaction Wheels, and 3
Torque Rods)
D. EPS – supplies monitored power to Attitude Determination and Attitude Control through
the flight computer.
2. Mechanical
A. Attitude Control is mounted internally within the satellite structure
B. The GPS and IMU are mounted internally within the satellite structure
51
C. The Sun Sensors are mounted externally on the satellite structures in opposite directions
acting along the same axis.
A list of the connectors used for interfacing can be seen below.
 Sun Sensor: DC Connector
 IMU: Accels - Connected to a circuit board.
 Gyro’s - RS422 Interface. Also mounted on circuit board.
 IMU computer – RS422/485, SPI, I2C
 GPS: RS422 Interface
 Reaction Wheels: RS422 Interface option.
 Torque Rod: Twist-pin micro-D style connector
Command and Data Handling
The purpose of C&DH is to ensure that data is properly transmitted between systems onboard the
satellite. This includes a method of storing data. During completion of the PDR, the C&DH
system design was narrowed down to non-radiation-hardened parts, an internal data storage
system, a high-performance processor, and COTS software. The design was further refined, as
required for the completion of the CDR. The components which compose the C&DH system can
be split into two categories, hardware and software.
Hardware
C&DH will utilize the Andrews Model 160 High Performance Flight Computer, which has the
following specifications:








100 MHz Processor
2GB Flash Memory
64MB SD RAM
Power Consumption (Nominal/Max): 5W/9W
Mass: 70 grams
Volume: 15x97x90 mm
Operating Temperature: -30 to +65o C
Input Power: 6.5V, 12V, 28V
Software
C&DH will purchase COTS software from Clyde Space, which will be modified to provide the
following modes of operation to the flight computer:
Startup:
Once the EPS has provided power to C&DH, C&DH will send a signal to OADC to begin
detumbling. Once successful detumbling has been verified by OADC, C&DH will signal EPS to
begin providing power to the payload and solar array deployment mechanisms. Upon verification
by Payload and EPS that deployment was successful, the flight computer will switch to standby
mode.
52
Standby:
In standby mode, C&DH will continuously receive requests from OADC to verify the satellite’s
orbit and attitude. This will be done via matrix calculations by the flight computer comparing the
current orbit with the desired orbit. The difference in orbits is then sent to OADC which will
correct the satellite’s course and orientation. OADC data from the IMU will be saved to the
flight computer’s flash memory after being converted using industry standard compression
algorithms.
C&DH will continuously receive, compress and record Solar Array Power Status, Solar Array
Position, Battery Charge/Discharge Status, and Power/Voltage to each subsystem component, as
received from EPS, to the flight computer’s flash memory. C&DH will send a true/false signal to
EPS to start/stop charging the battery when EPS’s sun sensor detects the satellite is in/out of the
sunlight. C&DH will also continuously provide EPS with commands to the gimbal to properly
orient the solar array.
C&DH will continuously receive, compress, and record temperature health data from TPS’s
seven temperature sensors to the flight computer’s flash memory. Consequently, C&DH will
signal EPS to stop power distribution to a particular component should that component’s
temperature exceed its operating range. This will minimize damage to the component should a
TPS failure occur.
C&DH will wait for a signal from payload for each MMOD impact reading. The particle’s
velocity vector relative to the satellite will then be numerically indexed alongside the satellite’s
current orientation. This information will be compressed and stored in the flight computer’s flash
memory. C&DH will then wait for a response from the COMM system signaling that the satellite
is within range of a ground station, at which point the flight computer will switch to downlink
mode.
Downlink:
C&DH will transfer all recorded data in the flight computer’s flash memory to the COMM
system, and continue transferring any newly recorded data while linked to the ground station.
C&DH will wait for ground commands, to be received through the COMM system, which will
be executed to the applicable subsystem. Such commands may include disabling of subsystem
components through EPS, or a command to initiate de-orbit, at which point the flight computer
will switch to de-orbit mode.
De-orbit:
C&DH will signal EPS to stop power distribution to all unnecessary subsystems, and signal
OADC to begin de-orbit procedures. After OADC signals that de-orbit procedures have been
completed, C&DH will signal EPS to stop power distribution to all subsystems.
53
Figure 39. Telemetry flow between C&DH interfaces.
Technology Readiness Level Advancement Plan
The flight hardware for the on-board computer is already at a TRL 8. The system has been flight
tested and will not be modified. The testing performed by the company prior to purchase of the
computer includes functional, vibration, mechanical shock, thermal cycling and thermal vacuum.
The flight software is at a TRL 6. This COTS software has been used previously, but slight
modifications will be made for this design. To bring the computer software up to TRL 7, inhouse DITL testing will be performed by BAI to detect any bugs, which will then be fixed.
Communications
The responsibility of the Communications Subsystem revolves around the satellite’s capability to
transmit data to the ground stations, receive commands, and successfully stay in contact with the
orbiting satellite. Without successful communication from the satellite to the ground stations, the
mission would be at severe risk. Loss of communications may result in a total loss of all control
over the satellite.
Design Outline
The final design utilizes a patch antenna along with a transmitter and receiver, which will all be
acquired from ISIS. The antenna, seen in figure 40, works on the s-band frequency and has an
opening angle of 85 degrees. The antenna has a right-hand circular polarization pattern and a
gain of 6 dB. A low antenna gain reduces the accuracy needed when directing the antenna
towards the ground stations for communication.
54
Figure 40. Solid model of S-band patch antenna
During downlink, the C&DH will send signals to the transmitter. The transmitter contains two
circuit boards: HI-RFE (Radio Front End), HI-CTR (Controller). The Controller Module is used
for whole data processing, data encoding, digital-to-analog (D/A) conversion of transmits
signals, transfer of control data to/from the on-board processor and power supply of both
modules. The RFE Module is the actual transmitter of the signal. Then the signal is sent to the
antenna which sends the data to GSS. The data sent to GSS includes telemetry data and payload
data.
The transmitter has a maximum data rate of 1 Mb/s, which exceeds the minimum data rate of
0.66 Mb/s at which data will be transmitted to the ground station. A similar model of the
transmitter to be acquired from ISIS can be seen in figure 41.
Figure 41. Surrey S-Band Transmitter
During uplink, GSS sends signals to the antenna at a maximum data rate of 1 Mb/s. The signal is
transferred from the antenna to the receiver. The receiver de-modulates and decodes the data sent
from GSS. These signals include emergency commands and OADC commands. A similar
receiver from Surrey can be seen in figure 42.
55
Figure 42. Surrey S-Band Receiver
A table of the component characteristics including power requirements, dimensions, mass and
temperature range requirement can be seen in table 6.
Table 6. COMM component characteristics
Frequency
Patch
Antenna
Transmitter
Receiver
Dimensions
(mm)
50 x 50 x 3.2
Mass (g)
S-band
Power
Requirement (W)
10
80
Temperature
Range (⁰C)
-120 to 120
S-band
S-band
5
5
95 x 46 x 15
150 x 180 x 30
100
130
-40 to 65
-40 to 45
Technology Readiness Level
The Technology Readiness Level, TRL, is inherently important because it can help to ensure that
the performance demands are met by the communication system. Choosing an initial TRL of 7
provides a system that has been adequately tested in relative environments. Although a TRL of 7
is still partially a prototype, the prototype may be developed further through funded testing to
reach a TRL of 8.
All three components (antenna, transmitter, and receiver) are at a TRL level of 7. They have
been tested in the operational environment. The proposed advancement plan to bring the TRL
level to 8 is further testing of the components.
Interfaces
The communications system contains interfaces with structures, C&DH, and GSS. C&DH sends
data via the transmitter which in turn directs it to the antenna. The antenna downlinks the data
package to the ground station. When uplinking, the data from the ground station is received by
the antenna which sends the data to the receiver. The receiver sends the data to be processed by
C&DH.
56
The electrical interfaces discussed consist of wiring. The three components have specific
connectors which allow them to transmit and receive the data. The antenna uses an SMA female
connector, the transmitter used a 12 way SMC connector and the receiver uses a TNC female
connector.
C&DH
Transmitter
Receiver
Antenna
GSS
Figure 43. COMM interface chart with GSS
The system also has a physical interface with the structure. All of the components are mounted
on to the structure via M4 fasteners, with the antenna being mounted on the outside. The chart
below details the outline of the operational process. The Communications Subsystem primarily
works with C&DH and GSS.
Safety and Mission Assurance
In order to assure the survival and success of the COMM system, testing will be done on several
aspects of the COMM system. Testing will ensure that the COMM system will satisfy all top
level requirements in the environment of space. Testing will also be done to ensure that the
COMM system will operate correctly despite experiencing forces and torques during rocket
launch and de-tumbling of the satellite.
Testing (to be done by supplier)
a.)
b.)
c.)
d.)
e.)
f.)
Functionality
Vibration
Mechanical Shock
Thermal Cycling
Thermal Vacuum
Total Ionizing Dose
57
Functional testing will be done to ensure the COMM system will operate as designed. The data
rate, frequency, and RF output power will be measured to ensure that they meet the top level
requirements.
Vibration testing will be done to ensure the COMM system will last the length of the mission.
Vibration testing will measure the fatigue life and resonant frequencies of the COMM system.
Mechanical shock testing will be done to ensure the COMM system will last through the launch
of the rocket and de-tumbling of the satellite. Shock testing will be the fragility or strength of the
COMM system while being loaded.
Thermal cycling testing, thermal vacuum testing and total ionizing dose (TID) testing will be
done to ensure the COMM system will last in the environment of space. Thermal cycling will
measure the COMM system’s ability to last while subjected to very high and very low
temperature experienced in space. Thermal vacuum testing will ensure the COMM system will
not experience a failure, such as outgassing, due to being in a vacuum. TID testing will ensure
the COMM system will not fail due to degradation from ionizing radiation.
Final Product
The integration of the seven subsystem designs described in Satellite Subsystem Design create
the final product. Figures 44a and 44b provide front and back views of the deployed satellite,
respectively. Figure 44c provides an internal view of the satellite. Figure 45 provides an
overview of the interfaces between the subsystems on BOSS.
Figure 44a. Front view of fully deployed satellite. A sun sensor and the front faces of the
DRAGONS and solar arrays are clearly visible.
58
Figure 44b. Back view of fully deployed satellite. Visible are the terminator tape deployables
and a sunsensor.
Figure 44c. Internal view of BOSS. All external components have been removed to provide
better visibility.
59
Figure 45. Interfaces found on BOSS
Operations
The process of transmitting data from the satellite to BAI facilities starts with the
communications system on the satellite. When in range of one of the selected ground stations,
data will be transmitted from the satellite to one of four NASA ground stations. Applicable
ground stations include Kennedy, Johnson, Whitesands, and Ames. STK based analysis
concluded that the satellite will provide a communications window at least twice a day. The
ground station(s) that the receive data for the particular day would forward it to BAI facilities. At
BAI facilities, the health data acquired will require about an hour of analysis per day to conclude
the satellite’s health status. As long as the health status is good, operations will continue
normally. If the health status is deemed bad, the relevant subsystem will be consulted for
analysis and troubleshooting.
An STK analysis determined that each ground station will have about 400 seconds of contact
with BOSS during the communications window. Since it was established that there will be a
communications window at least twice per day between, sending data at 1 Mb/s would yield a
maximum data transfer of at least 3200 Mb per day. The daily maximum amount of expected
data collected by the satellite is 528 Mb, so all stored data will be transferred from the satellite
daily.
The process of uplinking commands to the satellite from the BAI facilities works very similarly
to the downlink process. When commands need to be sent to the satellite, the commands are sent
to the applicable NASA ground station to uplink the command when the satellite is within in
60
range. Confirmation that the commands have been completed will be downlinked to the
applicable NASA ground station, and forwarded to BAI facilities. The confirmation data is
verified by the same engineer observing health data.
The data acquired from this mission will update current MMOD models, which is of great
interest to the DOD and NASA, as well as non-government aerospace agencies. Therefore, it was
negotiated that NASA ground stations will be used to communicate with BOSS at no cost to
BAI, and with the understanding that NASA and other agencies will have access to the space
debris data. Hence, the costs associated with GSS are purely labor related. The estimated cost of
labor to monitor data would be $27,920 for 698 days at an average salary of $40/hour.
Considerations were made for scenarios in which analysis and troubleshooting was necessary
due to bad health status, so additional funds were allocated for this contingency.
Risk
The risks associated with the BOSS subsystems are outlined in the risk analysis matrix. Each risk
has an associated criticality based on the severity and likelihood of the event to occur. A
preventative action is outlined for each risk to ensure the risk is not fatal to the mission. The risk
matrix for BOSS begins on page 62.
61
Failure Mode
COMM
Antenna Failure
Transmitter Error
Tracking Error
Severtity (15)
Liklihood (15)
Criticality
Failure Cause
Failure Effects
Detection
Debris Impact
Component Failure
5
1
5
COMM to GSS
Power loss
Loss of GSS COMM
4
2
8
COMM to GSS
Not in range of station Loss of GSS COMM
2
3
6
COMM to GSS
Preventative
Action
Position antennas
out of impact zone
Backup battery
Attitude/Orbit
Control
Attitude/Orbit
Control
Improper orientation
Unreliable contact with
GSS
1
3
3
COMM to GSS
Electromagnetic
Interference
Poor signal
2
5
10
COMM to GSS
Improper Data Rate
Poor signal
2
4
8
COMM to GSS
High noise
Loss of contact with
satellite
Poor signal
2
2
4
COMM to GSS
Loss of GSS COMM
4
1
4
COMM to GSS
Improper Timing
Loss of GSS COMM
2
3
6
COMM to GSS
Thermal Deterioration
Component Failure
3
2
6
COMM to GSS
Depris Imact
Component Failure
5
1
5
COMM to GSS
(Gyro) Drift
1. Decay of Gyros and
their performance over
time.
2. Could be severe in a
short amount of time
or minor over a long
period of time.
1
1
1
Software
Cross-check Data
V&V
(Gyro) Bias Stability
1. Changes over time
creating more noise in
signals.
2. Could produce more
work at end of life in
data analysis.
3. Could skew data
1
1
1
Software
Cross-check Data
V&V
Filtering
Durability Failure
Insulation
C&DH Data
Adjustment
System Reset/Filter
Attitude/Orbit
Control
Attitude/Orbit
Control
Proper Themal
Protection
Proper Positioning
OADC
IMU Failure
1. Could cause
temporary malfunction
(Gyro) Random Walk in the software and
Environmental
hardware.
Conditions
2. Causes switching of
bits
2
3
6
(Gyro) Jitters
1. Causes Excess
Vibration on unit
2. Could skew Data
1
2
2
Manufacturing Error
1. Contamination of
Unit
2. Loss or failure of
Unit
3. Loss of Gyro or Accel
5
3
15
Loss of power to one
Gyro
Loss of Gyro on one
axis
3
1
3
Software
Cross-check data
Feedback Control
Software
Cross-check Data
Cross-check data
Cross-check Data
Software
"No real
preventative Action as caused by
environment"
2. Program software
and Hardware to
expect Random
Error and recover.
V&V
V&V
Improved Quality
Assurance
V&V
Failure Mode
Failure Cause
Failure Effects
Severtity (1- Liklihood (15)
5)
Criticality
Detection
Preventative
Action
OADC
V&V Test of system
with loss of power.
2nd Redundant Sun
Sensor
3
Software
Cross-check Data
3
15
Software
Cross-check Data
V&V
Improved Quality
Assurance
3
1
3
Software
Cross-check Data
V&V Test of system
with loss of power
1. Failure of the part
before launch.
2. No GPS position data
5
3
15
Software
Cross-check Data
V&V
Improved Quality
Assurance
Sign Interference from
Torque Rods
1.Cause
electromagnetic
Interference
2. Skew all positional
data of system.
3
2
6
Hard to detect
V&V
Loss of Power
1. Loss of control on
one axis
2. Loss of control of
spacecraft
3
1
3
Software
Cross-check Data
V&V
Redundant 4th
Wheel
Manufacturing Error
1. Premature loss of
control on one axis
2. Loss of control of
spacecraft
5
3
15
Software
Cross-check Data
Quality Assurance
V&V
Redundant 4th
Wheel
Torque Rod Failure
Manufacturing Error
1. Loss of control of
spacecraft
5
3
15
De-orbit Module
Failure
Manufacturing Error
1. Spacecraft will not
de-orbit in 25 year
deadline
5
3
15
Loss of Power
No sun sensor data
3
Manufacturing/Vibe
Testing Error
1. Failure of the part
before launch.
2. No sun sensor data
5
Loss of Power
No relative positioning
data
Manufacturing Error
1
Sun Sensor Failure
GPS Failure
Attitude
Determination
(System)
Reaction Wheel
Failure
Cross-check Data
GPS altitude Data
V&V
Improved Quality
Assurance
Quality Assurance
V&V
2nd Module
Payload
Sensor Unit(s) Failure
Control Box Failure
Resistive grid
malfunctions
No data will be
produced from grid
Wiring from box to unit Unable to store or
becomes destroyed
process data from unit
5
5
1
3
5
15
Loss of data from
particular sensor to
electronics box
Environmental
testing (thermal &
vibration). Careful
placement of
computer by TPS.
No sensor data in
downlink
Environmental
testing (thermal &
vibration). Use of
space grade wires
carefully placed
where no impacts
will occur.
63
Failure Mode
Failure Cause
Failure Effects
Detachment latch
malfunctions
DRAGONS Sensor is
rendered useless
Severtity (1- Liklihood (15)
5)
Criticality
5
2
Detection
Preventative
Action
10
Switches placed on Verify with
mechanical stop
functional test
4
Verify with
functional test and
Switches placed on meneuver satellite to
mechanical stop
ensure deployment
failure, not Sensor
failure.
Deployment Failure
Locking malfunction
DRAGONS Sensor is
not secure, data may
become skewed
Temperature range is
exceeded
No data will be
transmitted between
systems
5
2
10
Environmental
testing (thermal &
Sudden lack of any
vibration). Careful
data
placement of
computer by TPS.
No payload data
5
1
5
Data shows zero hits
Failure to begin
detumbling
5
1
5
OADC Data would DITL Testing prior
show
to flight
Failure to release
EPS/sensors
5
1
5
Satellite would stop DITL Testing prior
transmitting
to flight
No data is transmittedno debris data gathered
5
1
5
No release of data
1
2
2
Software/Data from
Short Mission Life
C&DH
5
1
5
Ensure Mechanism
Software/Data from
is purchased from a
C&DH
reputable source
4
2
8
Software/Data from
Solar Cell Testing
C&DH
2
1
2
Software
2
2
C&DH
Computer Hardware
Failure
Software error in
algorithm
Hardware Error
Programming Error
Excessive vibrations
during launch
DITL Testing prior
to flight
Vibration testing
prior to flight.
EPS
Battery Damage/Loss
Degradation over time of Battery storage
capacity
Battery Failure
Safety Mechanism
Relay Fails
Solar Cell Failure
Cover Glass Failure
Solar Array Failure
1. Batter over
discharges
2.
Battery over chargers
Loss of Capable Power
Production
Individual Solar Cell
Failure
Cover Glass testing
Testing of
Software/Data from
Deployment
C&DH
Methods
Software/Data from Simulations of
C&DH
Impacts
Deployment Failure
Loss of Capable Power
Production
4
2
8
Debris Impact
Solar Cell
Damage/Failure
1
4
4
Loss of Power to
Motor
Inability to Directe
Solar Arrays at Sun
4
2
8
Software/Data from Simulations with
C&DH
Loss of Power
Loss of Range of
Motion
Reduced Ability to
Direct Solar Arrays at
the Sun
2
2
4
Software/Data from Simulations with
C&DH
Partial Power Loss
4
2
8
Software/Data from
Back up Electronics
C&DH
Gimbal Failure
Failed
Electronics/Interconnec
Electronics/Interconnec Component Failure
ts
ts
64
Failure Mode
Severtity (1- Liklihood (15)
5)
Criticality
Detection
Preventative
Action
Failure Cause
Failure Effects
Large Debris Impact
Satellite Failure and
Mission Failure
5
1
5
Excess loading
Satellite Failure and
mission failure
5
2
10
Panel (2) attachment
failure
Unable to generate
energy for satellites
components
5
2
10
Loss of power
Environmental
testing (thermal &
vibration).
Dragon sensor
attachment failure
Unable to collect data
on debris impact
5
2
10
Unable to generate
debris impact data
Environmental
testing (thermal &
vibration).
Antenna (2) attachment Unable to communicate
failure
with satellite
5
2
10
Inability to
communicate with
satellite
Environmental
testing (thermal &
vibration).
Thermal mirror
attachment failure
5
2
10
Environmental
Overheating causing
testing (thermal &
multiple failures
vibration).
Attitude sensor & IMU
Unable to maintain orbit
attachment failure
5
2
10
Loss of orbit
Environmental
testing (thermal &
vibration).
CD&H attachment
failure
Unable to record and
store data
5
2
10
Loss of data
communication
Environmental
testing (thermal &
vibration).
Transmitter and
receiver attachment
failure
Unable to communicate
with satellite
5
2
10
Loss of
Environmental
communication with testing (thermal &
satellite
vibration).
Damage from Debris
Components exposed
to harsh space
environment
4
1
4
.Choose strong
Components are not
insulation material or
operational under
incorporate backup
certain temperatures
thermal protection
Absorptivity of heat
increases and
reflectivity decreases
4
3
12
Components are not
Protect the silver
operational under
with nichrome
certain temperatures
2
Allowing testing to
run for longer
Components such as
periods of time to
Power system seem
properly vacuum
to underperform
and running multiple
tests.
2
Testing of assembled
Components such as satellite. Multiple
Power system seem loops of pipe as
to underperform
opposed to one
large pipe.
Structure
Frame Structure
Failure
Deployment
attachment failure
Frame attachment
failure
Internal attachment
housing failure
Unable to reflect heat
from satellite
Environmental
Complete failure of
testing (thermal &
satellite
vibration).
Environmental
Complete failure of
testing (thermal &
satellite
vibration).
TPS
Break in Multilayer
insulation
Exposed to space
Aluminum Optical
environment for long
Solar Reflector Oxides
periods of time
Temperature and
Vacuum space
Environment Testing
Error
Thermistors stop
recording data
Lack of accurate
testing due to false
simulation of
environment by either
failing to test thermal
cycles or not reaching
perfect vacuum
Thermal Protection
System will
underperform
Failed electronics/
wiring
Thermal protection
temperature will be
collected
1
2
2
1
65
Technology Readiness Level
The technology readiness level matrix seen below detailed the overall TRL of the current components. A plan
for advancement has been outlined for applicable components in their respective sections.
Terminator Tape
Function
Appropriate Scale
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
Overall TRL
Fit
X
X
Form
Unit Description
Launch Operation
X
Space Environment
X
Relevant Environment
Laboratory Environment
Prototype
Developmental Model
Brassboard
Flight Qualified
Payload
DRAGONS Sensor
Hinges Spring Loaded
Hold Down Release Mechanism
Software
C&DH
On-board Computer
Software
COMMS
Antenna
Transmitter
Receiver
EPS
Solar Arrays
Power Conditioning Unit
Battery
Solar Array Gimbal
Space Rated Electrical Wiring
Distribution Bus
Interfaces
OADC
GPS
IMU
Sun Sensor
Reaction Wheels
Torque Rods
Breadboard
Concept
X
Technology Readiness and Technical Performance Matrix
Red = Below TRL 3
Demonstration Units
Environment
Yellow = TRL 3, 4 &
5
Green = TRL 6 and
Above
White = Unknown
Exists
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
66
Structure
Mechanical Components
Frame
Side Panels
Software Housing
Reflective Mirror Support
Attitude Sensor Housing/Support
DRAGONS box fixtures
Antenna Attachment/Support
Transmitter and Receiver Attachment
Mechanisms
Solar Panel Deployment Mechanism
Payload Deployment Mechanism
TPS
Multi-Layer Insulation
Heat Pipes
Solar Reflectors
GSS
BAI Facilities
NASA Ground Stations
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
67
Overall TRL
Appropriate Scale
Function
Fit
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
Form
Unit Description
Launch Operation
Space Environment
Relevant Environment
Laboratory Environment
Flight Qualified
Prototype
Brassboard
Breadboard
Concept
X Exists
Developmental Model
Red = Below TRL 3
Yellow = TRL 3, 4 & 5
Green = TRL 6 and Above
White = Unknown
Technology Readiness and Technical Performance Matrix
Demonstration Units
Environment
Verification and Validation
The verification of requirements is needed in order to ensure that all requirements are met for a successful mission. The requirements verification
matrix below shows the requirements for the BOSS mission. Each requirement can be traced to the originating document and clearly states how it is
to be verified.
Verification Matrix
Requirement
No.
05.0
Payload
05.0
Document
Shall Statement
Verification Success
Criteria
MDR
Payload
instruments and
sensors shall
provide sufficient
capabilities for
debris detection
method.
The BOSS payload
shall utilize 6
DRAGONS units
to detect location
and particle
characteristics of
small space debris
in the near-Earth
environment.
Each DRAGONS
unit shall have a
maximum power
consumption of 810 W.
The BOSS payload
shall utilize
framework to
structurally
arrange
DRAGONS units.
05.1
PDR Trade
Study
05.1.1
DRAGONS
Specs
05.2
DRAGONS
Specs
Verification
Method
Facility or
Lab
Phase
Acceptance
Requirement
Preflight
Acceptance
Performing
Organization
Sensors provide
Test
sufficient data
when impacted by a
particle.
Bravo
Industries
Testing
Lab
D
Yes
Yes
Bravo
Aerospace
Industries
Sensors provide
location and energy
data after impact
with particle.
Bravo
Industries
Testing
Lab
D
Yes
Yes
Bravo
Aerospace
Industries
Bravo
Industries
Testing
Lab
C
Yes
No
Bravo
Aerospace
Industries
Bravo
Industries
Testing
Lab
C
Yes
No
Bravo
Aerospace
Industries
Test
The power required Test
for operation of a
DRAGONS unit is
between 8 and 10
W.
The structure is
Analysis
strong enough to
securely hold the
DRAGONS units in
place in a space
environment.
05.3
PDR Trade
Study
05.4
PDR Trade
Study
The BOSS payload
shall be folded
compactly around
the satellite frame
during vehicle
launch.
The BOSS payload
shall utilize a
deployment
mechanism from
the satellite main
framework in
order to deploy the
sensor array
securely.
The payload fits
around the 0.6 m x
0.6 m x 0.6 m
securely without
protruding.
Inspection
Bravo
Industries
Testing
Lab
C
Yes
Yes
Bravo
Aerospace
Industries
The sensor array
deploys in space
environment
(vacuum) securely,
with acceptable
vibration and
dampening time.
Test
Bravo
Industries
Testing
Lab
C
Yes
Yes
Bravo
Aerospace
Industries
C&DH shall
successfully
collect and process
data to be
transmitted to the
ground station.
The C&DH shall
transmit operations
data to the
COMMS.
The C&DH shall
send command
data to the OADC.
Data is collected
and processed
without error.
Test
Bravo
Industries
Testing
Lab
C
Yes
Yes
Bravo
Aerospace
Industries
Data is transmitted
without error.
Test
C
Yes
Yes
Bravo
Aerospace
Industries
Command data is
transmitted without
error.
Test
C
Yes
Yes
Bravo
Aerospace
Industries
The C&DH shall
send command
data to the
payload.
Command data is
transmitted without
error.
Test
Bravo
Industries
Testing
Lab
Bravo
Industries
Testing
Lab
Bravo
Industries
Testing
Lab
C
Yes
Yes
Bravo
Aerospace
Industries
06.1 C&DH
06.1.0
CDH MDR
06.1.1
CDH MDR
06.1.2
CDH MDR
06.1.3
CDHMDR
69
06.2
COMM
06.2.0
06.2.1
06.2.2
06.2.3
COMM
The COMM shall
Requirements successfully
transmit “energy
data” to ground
station.
COMM
The COMM shall
Requirements be capable of
uploading data at a
range greater than
700 km.
COMMRequi The COMM shall
rements
utilize a gain
greater than 0db.
COMM
The COMM shall
Requirements have a minimum
data rate of 1
Mb/s.
06.2.4
COMM
The COMM shall
Requirements utilize less than 15
W of power.
06.3 EPS
06.3.0
EPSMDR.pd
f
The EPS shall
store and distribute
adequate power
necessary for
operation of all
electrical
components at
peak load.
Data is transmitted
without error.
Test
Bravo
Industries
Testing
Lab
C
Yes
Yes
Bravo
Aerospace
Industries
Minimum range of
700 km for data
transfer is satisfied.
Analysis
Bravo
Industries
Testing
Lab
C
Yes
No
Bravo
Aerospace
Industries
Satellite is able to
successfully
transmit readable
data to the ground
station.
The data can be
sent at a rate of 1
Mb/s
Inspection
Bravo
Industries
Testing
Lab
D
Yes
No
Bravo
Aerospace
Industries
Analysis
Bravo
Industries
Testing
Lab
D
Yes
No
Bravo
Aerospace
Industries
The required power
for operation is
under 15 W.
Analysis,
Test
Bravo
Industries
Testing
Lab
D
Yes
No
Bravo
Aerospace
Industries
Peak power load
can be provided
from battery and
charged during sun
exposure.
Test
Bravo
Industries
Testing
Lab
C
Yes
Yes
Bravo
Aerospace
Industries
70
06.3.1
06.3.2
06.3.3
06.3.4
06.4 OADC
06.4.0.1
EPS MDR
The EPS shall
provide power for
the satellite for the
duration of the
three year mission.
EPS
The EPS shall
Requirements generate power
from the
environment.
EPS
The EPS shall
Requirements provide power to
the components
while not exposed
to the Sun.
EPS
The EPS shall
Requirements have a distribution
bus to distribute
adequate power to
each subsystem.
OADC
The OADC shall
Requirements determine the
relative position of
the spacecraft at
impact.
Solar panels and
battery can same
power at the end of
the 3 year mission.
Test
Manufactur C
er Facility
Yes
No
Component
Manufactur
er
Satellite can
maintain power
while in orbit,
without carrying
fuel.
Power levels are
maintained during
the time period of
no sun exposure
Test
Bravo
Industries
Testing
Lab
C
Yes
No
Bravo
Aerospace
Industries
Test
Bravo
Industries
Testing
Lab
C
Yes
Yes
Bravo
Aerospace
Industries
The bus distributes
power that each
subsystem requests
Test
Bravo
Industries
Testing
Lab
D
Yes
Yes
Bravo
Aerospace
Industries
1. Position data
received by GSS
2. Control
commands produce
expected attitude
changes.
Test
NASA
Testing
Facility
D
Yes
Yes
NASA
71
06.4.0.2
OADC
The OADC shall
Requirements determine motion
relative to the
Earth at impact.
06.4.0.3
OADC
The OADC shall
Requirements transmit positional
1. Attitude rate data
received from
satellite and meets
expected motion.
Test
NASA
Testing
Facility
D
Yes
Yes
NASA
Data successfully
transmitted.
Test
NASA
Testing
Facility
D
Yes
Yes
NASA
1. Position data
received by GSS
Test
Facility
Last
Yes
Yes
Bravo
Industries
1. GPS signals are
initiated; seen from
GSS
1. GPS operates
correctly within the
voltage range
Test
Facility
Last
Yes
Yes
NASA
Inspection
Facility
D
Yes
Yes
NASA
1. GPS receives
sufficient power to
operate correctly
for the input
voltage
1. GPS signals are
received every 20
milliseconds
Inspection
Facility
D
Yes
Yes
NASA
Inspection
Facility
D
Yes
Yes
NASA
1. GPS remains
operational within
given temperature
range
Inspection
Facility
D
Yes
Yes
NASA
data to the C&DH
at impact.
6.4.1 GPS
6.4.1.0
ICD
6.4.1.1
ICD
6.4.1.2
GPS Spec
Sheet
6.4.1.3
GPS Spec
Sheet
ICD
6.4.1.4
GPS Spec
Sheet
6.4.1.5
GPS Spec
Sheet
GPS position data
shall be
transmitted to
C&DH
GPS shall receive
commands from
C&DH
GPS shall operate
for input voltage
between 18 and 38
V
GPS shall receive
5.5 W at 28 V
unregulated from
C&DH
GPS shall update
position with a
frequency of 50
Hz.
GPS shall operate
between -20° and
+50° C
72
6.4.2 IMU
6.4.2.0
CDR
6.4.2.1
CDR
6.4.2.2
ICD
6.4.2.3
CDR
IMU shall process
accelerometer and
gyroscope
measurements
1. Filtered data
received, bias and
noise levels within
tolerance of
filtering
IMU shall
1. Error in
determine relative orientation does not
satellite orientation exceed 0.5°
within 0.5°
Filtered IMU
1. Data is received
angular rates shall from satellite
be transmitted to
C&DH
Satellite position
1. Data is received
shall be
from satellite
transmitted to
C&DH
Test
Facility
D
Yes
Yes
NASA
Test
Facility
D
Yes
Yes
NASA
Test
Facility
D
Yes
Yes
NASA
Test
Facility
D
Yes
Yes
NASA
Sun sensor
position data shall
be transmitted to
C&DH
Sun sensor shall
receive commands
from C&DH
Sun sensor shall
determine the
position relative to
the sun within 1°
Sun sensor shall
operate at ±12 V
1. Sun sensor
position data
received by GSS
after downlink
1. Sun sensors turn
on after initial
command.
1. Error does not
exceed 1°
Test
Facility
Last
Yes
Yes
NASA
Test
Facility
Last
Yes
Yes
NASA
Test
Facility
D
Yes
Yes
NASA
1. Sun sensor
operates at voltage
Test
Facility
D
Yes
Yes
NASA
6.4.3 Sun
Sensor
6.4.3.0
ICD
6.4.3.1
ICD
6.4.3.2
Sun Sensor
Spec Sheet
6.4.3.3
Sun Sensor
Spec Sheet
73
6.4.3.4
Sun Sensor
Spec Sheet
Sun sensor shall
use a maximum of
0.1 W
6.4.3.5
Sun Sensor
Spec Sheet
Sun sensor shall
operate between 20° and +50° C
6.4.4
Control
6.4.4.0
MCR
6.4.4.1
CDR
6.4.4.2
MCR
1. Sun sensors are
operating within
acceptable power
range
1. Sun sensor
operates between
temperature
boundaries
Inspection
Facility
Last
Yes
Yes
NASA
Test
Facility
D
Yes
Yes
NASA
Spacecraft shall
maintain and
adjust attitude
autonomously
Position and rate
data logs show
autonomous
reaction to
changing attitude
Test
Facility
Last
Yes
Yes
NASA
Satellite shall
maintain an
autonomous
orientation within
5° from desired
Spacecraft shall
maintain and
adjust attitude as
commanded
Measured deviation
from desired
orientation in any
axis is below 5°
Inspection
Facility
D
Yes
Yes
NASA
Spacecraft responds Test
to commands and
reorients itself to
given command
Facility
Last
Yes
Yes
NASA
Reaction wheels
shall receive
commands from
C&DH
Control commands
produce expected
attitude changes
Facility
Last
Yes
Yes
NASA
6.4.5
Reaction
Wheels
6.4.5.0
ICD
Test
74
6.4.5.1
ICD
Reaction wheels
shall send angular
velocity data to
C&DH
6.4.5.2
ICD
Reaction wheels
shall operate
between 1.8 and
10 W
6.4.5.3
Reaction
Wheel Spec
Sheet
6.4.5.4
6.4.5.5
Reaction wheel
angular velocity
data received by
GSS after downlink
Test
Facility
Last
Yes
Yes
NASA
1. Reaction wheels
draw power in
expected range
2. Reaction wheels
remain in standby
at 1.8 W
Reaction wheels
1. Reaction wheels
shall operate at 17- operate properly for
35 V DC
given voltage range
Test
Facility
Last
Yes
Yes
NASA
Inspection
Facility
D
Yes
Yes
CDR
Reaction wheels
shall provide
orientation control
within 0.5°
Inspection
Facility
D
Yes
Yes
NASA
Reaction
Wheel Spec
Sheet
Reaction wheels
shall operate
between -20° and
+50° C
Inspection
Facility
D
Yes
Yes
NASA
1. Measured
deviation from
commanded
orientation in any
axis is below 0.5°
1. Reaction wheels
operate properly in
given temperature
range
Yes
6.4.6
Torque
Rods
6.4.6.0
ICD
6.4.6.1
ICD
Torque rods shall
receive commands
from C&DH
Torque rods shall
receive power
regulated by
C&DH
1. Torque rods turn
on after given
command
1. Torque rods
receive power
necessary during
momentum dumps
Test
Facility
Last
Yes
Yes
NASA
Test
Facility
Last
Yes
Yes
NASA
75
6.4.7
Terminator
Tape
CDR
Terminator tape
shall deploy upon
command
1. Terminator tape
deploys when given
command at end of
life
Test
Facility
Last
Yes
Yes
NASA
06.5.0.1
Structure
MDR
The structure has
the adequate
strength to survive
in the space
environment for
three years.
Analysis
Bravo
Industries
Analysis
Lab 1
D
Yes
No
Bravo
Aerospace
Industries
06.5.0.2
TPS MDR
The structure shall
maintain integrity
and support
subsystems
throughout the
three year mission
life.
The TPS shall
maintain satellite
and subsystems
temperatures
within ranges of
normal operation.
Subsystems remain
at constant
temperatures
whether or not it’s
in the Earth's
shadow
Test
Bravo
Industries
Testing
Lab
D
Yes
No
Bravo
Aerospace
Industries
06.5.1
Structure
The satellite
Requirements structure and
The satellite frame
and structure
remain intact after
launch. Also, all
components in the
interior of the
satellite that are
fastened to the
satellite remain
fastened after
launch.
Test
Bravo
Industries
Testing
Lab
D
Yes
No
Bravo
Aerospace
Industries
6.4.7.0
06.5
Structure/
TPS
structural
attachments shall
withstand all
forces and
vibrations endured
during launch.
76
06.5.2
06.5.3
06.5.4
06.5.5
07.0
Mission
Operations
07.0
Structure
The structure shall
Requirements have all
deployment
systems stowed
during launch.
TPS
The satellite shall
Requirements be protected from
the space
environment
(radiation)
throughout the
three year mission
life.
TPS
The TPS uses
Requirements copper heat pipes
filled with
ammonia to
transfer heat
between satellite
components.
TPS
The TPS shall
Requirements transmit
temperature data to
the C&DH.
Top Level
The Mission
Requirements Operations shall
generate a plan
which allows for
the success of the
mission
requirements.
All deployment
systems are stowed
securely.
Test
Bravo
Industries
Testing
Lab
D
Yes
Yes
Bravo
Aerospace
Industries
Thermal insulation
provides adequate
shield from
radiation in the
environment.
Analysis
Bravo
Industries
Analysis
Lab
D
Yes
No
Bravo
Aerospace
Industries
Heat pipes must not
deform, degrade, or
lose physical
properties when put
under stress of
environment
Test
Bravo
Industries
Testing
Lab
D
Yes
Yes
Bravo
Aerospace
Industries
Data is successfully
transmitted.
Test
Bravo
Industries
Facility
D
Yes
Yes
Bravo
Aerospace
Industries
Plan is successfully
implemented.
Inspection
Bravo
Industries
Facility
C
No
No
Bravo
Aerospace
Industries
77
08.0
Launch
Vehicle
08.0
Top Level
The launch vehicle
Requirements shall safely deliver
BOSS to a 700 km
polar orbit.
Launch vehicle
safely delivers the
satellite to specified
orbit.
Analysis,
Test
Facility
D
Yes
Yes
Space
Exploration
Technologies
Corporation
A ground station
can communicate
with the satellite at
least once per day.
Analysis
Facility
C
Yes
No
Bravo
Aerospace
Industries
Receive and review
the health data for
an hour each day.
Inspection
Facility
D
Yes
No
Bravo
Aerospace
Industries
Facility
D
Yes
No
Bravo
Aerospace
Industries
09.0 GSS
09.0
09.1
09.2
GSS
The GSS shall be
Requirements able to downlink
and uplink data to
the satellite at least
once per day.
GSS
The GSS shall
Requirements monitor the health
information
(power,
temperature,
electrical
component status)
of the satellite and
payload for one
hour per day.
GSS
The GSS shall be
Requirements able to send
commands to
ground stations for
uplink to satellite
at least once per
day.
Have ability to send Test
commands to
satellite once a day
78
09.3
GSS
At least one
Requirements NASA ground
station shall be
able to receive
data from the
satellite at least
once per day.
Satellite is in range
and satellite
antenna downlinks
data to NASA
ground stations at
least once a day
Analysis
Facility
D
Yes
No
Bravo
Aerospace
Industries
The validation table below ensures that proper testing is done and validated by the customer, NASA. The validation table states the validation
parameters based on the mission’s ConOps and how each parameter is to be validated.
Validation
Product #
5.0
Payload
5.0
Activity
Objective
Validation
Method
Facility or
Lab
Phase
Performing
Organization
NASA/ DRAGONS developers
verify capability of three
detections technologies function
Ensure payload systems
detect impacts accurately
and individual detections
systems are compatible
with each other
Tests
C
Bravo
Aerospace
Industries/
USNA
5.1
NASA will verify six DRAGONS Ensure six DRAGONS unit Expert
unit layout consistent with CDR
present, functioning, and in Inspection
design technical drawings
an acceptable debris
detection layout
Two-stage
light gas
gun facility
of the
University
of Kent
Bravo
Industries
Testing Lab
C
Bravo
Aerospace
Industries/
NASA
5.2
NASA validate energy
consumption of individual and
combination of all DRAGONS
unit(s) system
To ensure each
DRAGONS unit is
consuming at max 8-10 W
of power
Test
Bravo
Industries
Testing Lab
C
Bravo
Aerospace
Industries/
NASA
5.3
NASA validate rigid and
acceptable structure to arrange
and contain DRAGONS units
throughout mission
Confirmed designed
Structure will be suitable
for mission requirements
and lifetime
Test (
Vibrations
and Load
testing)
Bravo
C
Industries
Testing Lab/
Vibrational
Bravo
Aerospace
Industries/
NASA/USN
79
Analysis
Lab/Compa
ny
A
5.4
NASA validates that’s
DRAGONS compactly fold onto
satellite structure.
Payload system compactly
fits around 0.5 x 0.6x 0.6
m^3 satellite structure
Expert
Inspection
Bravo
Industries
Testing Lab
C
Bravo
Aerospace
Industries
5.5
NASA validates that payload
system deploys fully and
communicates via micro switches
of successful deployment.
Ensure system fully
deploys for accurate
measurements and that
GSS will confirm
deployment during mission
Test
Bravo
Industries
Testing Lab
C
Bravo
Aerospace
Industries
Customer will evaluate computer
software
1. Ensure desired modes
are achieved
DITL
Testing
Bravo
Industries
Testind Lab
A
NASA
Customer will evaluate the energy Ensure quality of "energy
data transferred to the ground
data" is acceptable
Test
Bravo
Industries
Testing Lab
C
Bravo
Aerospace
Industries
6.2.1
Customer will evaluate the
maximum range of the COMM
system
Ensure range of the
COMM system is
acceptable
Analysis
Bravo
Industries
Testing Lab
C
Bravo
Aerospace
Industries
6.2.2
Customer will evaluate the gain
of the antenna
Ensure the gain of the
antenna is acceptable
Inspection
Bravo
Industries
Testing Lab
C
Bravo
Aerospace
Industries
6.1
C&DH
6.1
6.2
COMM
6.2.0
80
6.2.3
Customer will evaluate the data
rate of the COMM system
Ensure the data rate is
acceptable
Analysis
Bravo
Industries
Testing Lab
C
Bravo
Aerospace
Industries
6.2.4
Customer will evaluate the power
required of the COMM system
Ensure the power required
is acceptable
Analysis,
Test
Bravo
Industries
Testing Lab
C
Bravo
Aerospace
Industries
Customer will evaluate the
storage capacity of the battery
Ensure battery capacity is
large enough for a peak
power load from
components
Battery
Storage
Test
Bravo
Industries
Testing Lab
C
Bravo
Aerospace
Industries
6.3.1
Customer will evaluate the life of
the EPS subsystem components
Customer will evaluate the ability
of EPS to collect power from the
environment
Bravo
Industries
Testing Lab
Bravo
Industries
Testing Lab
C
6.3.2
Ensure all components will Degradalast the length of the
tion Tests
mission
Ensure the satellite can
Test
maintain power without the
need to carry fuel
Bravo
Aerospace
Industries
Bravo
Aerospace
Industries
6.3.3
Customer will evaluate the
storage capacity of the battery
and length of earth eclipse
Customer will evaluate the
success of EPS's distribution bus
and voltage regulation
Ensure the battery isn't
over-discharged during
eclipses
Ensure the bus distributes
correct power to different
components
Ammeter
Test
Bravo
Industries
Testing Lab
Bravo
Industries
Testing Lab
C
Customer will evaluate the wiring
used with the electrical
components
Ensure the wiring is space
grade and will withstand
space environment
Vacuum
Test
Bravo
Industries
Testing Lab
C
6.3 EPS
6.3.0
6.3.4
6.3.5
Ammeter
Test
C
C
Bravo
Aerospace
Industries
Bravo
Aerospace
Industries
Bravo
Aerospace
Industries
OADC
81
6.4.0
Customer will evaluate spacecraft
relative position determination
1. Determination system is
operating properly
2. Relative position data is
accurate
Test
Facility
D
NASA
6.4.1
Customer will evaluate spacecraft
relative motion determination
Determination system is
operating properly
2. Relative motion data is
accurate
Test
Facility
D
NASA
Check flight computer data
Flight computer is
receiving position data
Test
Facility
C
Bravo
Industries
6.4.2.2
Check operating voltage of the
GPS
Verify if the operating
range is acceptable.
Test
Facility
C
Bravo
Industries
6.4.2.3
Determine power draw of the
GPS
Verify GPS receives
approximately 5.5W at 28
V unregulated
Test
Facility
C
Bravo
Industries
6.4.2.4
Test and refine GPS refresh rate.
GPS updates position at the Test
appropriate frequency
Facility
C
Bravo
Industries
6.4.2.5
Thermal test of GPS
Verify that GPS operates
within the thermal
boundaries
Test
Lab
C
Bravo
Industries
Verify the IMU flight
computer correctly
processes the
accelerometer and
gyroscope measurements
Test
University
C
University
2. Bravo
Industries
Verify final attitude output
from the IMU is accurate
to within 0.5 degrees
Test
University
C
University
2. Bravo
Industries
6.4.2 GPS
6.4.2.1
6.4.3 IMU
6.4.3.0
Check filtered angular rates from
IMU
6.4.3.1
Check position output from IMU
82
6.4.3.2
View data sent to C&DH
Flight computer receives
filtered angular rates
2. Flight computer receives
position data
Test
Facility
C
Bravo
Industries
6.4.4 Sun
Sensor
6.4.4.0
Check flight computer data
Flight computer receives
sun sensor position data
Test
Facility
C
Bravo
Industries
6.4.4.1
Check sun sensor commands
from C&DH
Verify that sun sensor
receives inputs from the
flight computer.
Test
Facility
C
Bravo
Industries
6.4.4.2
Check position output of sun
sensors
Sun sensor position data is
within 1 degree
Test
Facility
C
Bravo
Industries
6.4.4.3
Check operating voltage of sun
sensors
Sun sensors operate for an
acceptable voltage
Test
Facility
C
Bravo
Industries
6.4.4.4
Check power draw of sun sensors
Sun sensors draw
acceptable power
Inspection
Facility
C
Bravo
Industries
6.4.4.5
Thermal test of sun sensors
Verify that the sun sensors
operate within the thermal
boundaries
Test
Lab
C
Bravo
Industries
6.4.5
Control
6.4.5.0
Analyze satellite attitude data
Verify autonomous control Test
is achieved using data
provided by flight
computer and control
system
Verify active (commanded) Test
control is achieved using
data provided by flight
computer and control
system
Facility
D
Bravo
Industries
Facility
D
Bravo
Industries
6.4.5.1
Analyze satellite attitude data
after commands
83
6.4.6
Reaction
Wheels
6.4.6.0
Check reaction wheels' reaction
to commands
Verify commands are
received from by the flight
computer
Test
Facility
C
Bravo
Industries
6.4.6.1
Check flight computer data
Test
Inspection
Facility
C
Bravo
Industries
6.4.6.2
Check power draw of reaction
wheels
Verify flight computer
receives angular velocity
data from the reaction
wheels
Verify reaction wheels
receive the appropriate
power
Test
Inspection
Facility
C
Bravo
Industries
6.4.6.3
Check operating voltage of
reaction wheels
Verify reaction wheels
operate within advertised
voltage range
Test
Inspection
Facility
C
Bravo
Industries
6.4.6.4
Thermal test of reaction wheels
Verify that the reaction
operate within the thermal
boundaries
Inspection
Facility
C
Bravo
Industries
6.4.7
Torque
Rods
6.4.7.0
Send commands to torque rods
Verify commands received
from the flight computer
Test
Facility
C
Bravo
Industries
Determine power draw of the
torque rods
Power draw of torque rods
is within acceptable ranges
Test
Facility
C
Bravo
Industries
Check commands sent by flight
computer
Review that the flight
computer sends command
to deploy the terminator
tape
Test
Facility
C
Bravo
Industries
6.4.7.1
6.4.8
Terminator Tape
6.4.8.0
84
6.4.8.1
6.5.0
Structure
6.5.0.1
6.5.0.2
6.5.1 TPS
6.5.1.1
6.5.1.2
6.5.1.3
6.5.1.4
6.5.1.5
Deployment of the terminator
tape
Terminator tape receives
command and deploys
Inspection
N/A
C
NASA
Customer will evaluate the
response of the structure to
launch conditions (loads and
associated vibrations).
Customer will evaluate the
response of the structure and its
components to high heat
conditions
Ensure the structure can
survive launch intact.
Analysis
and then
Test
Bravo
Industries
Testing Lab
C
Bravo
Aerospace
Industries
Ensure that the structure
will burn up in de-orbit.
Analysis
Bravo
Industries
Testing Lab
C
Bravo
Aerospace
Industries
Customer will evaluate that the
TPS is operational passively with
little electrical consumption
Use of battery and
electrical wiring only
needed for temperature
sensors
Battery
Test
Facility
C
Bravo
Aerospace
Industries
Customer will evaluate the MLI
that protects a satellite with the
dimensions of 0.6m x 0.6m x
0.5m from radiation
Customer will evaluate the optical
solar reflectors to reduce
absorptivity
Customer will evaluate the
temperature sensors within the
satellite
Subsystems remain at
constant temperatures
whether or not it’s in the
Earth's shadow
Reflectivity shall be larger
than 90%
Radiative
Heat
Transfer
Test
Spectropho
tometry
Test
Environme
ntal Test
Chambers
Lab
C
Bravo
Aerospace
Industries
Lab
C
Lab
C
Bravo
Aerospace
Industries
Bravo
Aerospace
Industries
Customer will evaluate the
durability of the TPS
The TPS will be functional
for two years in the harsh
space environment
Testing
Lab
C
The satellite must be kept
between -10०C and 30०C
Bravo
Aerospace
Industries
85
System Performance Requirements Flow down Chart
Release Latch
Error
Sensor
Deployment
Incomplete
Deployment
Thermal
Deformation
Hinge Damage
Switch Activation
Payload
Allocation Error
Powe/Data
Unregistered
Particle Hit
Impact Error
Hardware
C&DH
Recieve data
Ricochet Off
Strucutre
Software
Transmit data
Software
System Performance
Requirements
Uplink
Emergancy
Commands
Size/Mass
Constraints
Hardware
Receiver
Capabilities
COMM
SNR
Hardware
Transmission
Range
Downlink
Data Rate
Software
Transmission
Quality
Orientation Error
Power Generation
Solar Panels
Exposure Error
Time of Orbit in
Sun Error
Solar Panel Sizing
Error
Battery
Overcharged
EPS
Power Storage
Storage
Degredation
Battery Over
Distcharged
Degradation Error
Thermal Failure
Power Distribution
OADC Error
C&DH Error
Voltage and
Current
28V Distribution
Payload Error
COMMS Error
System Performance Requirements chart is continued on the next page.
GPS
Positioning
Sun Sensor
Sunsensors
Rate
IMU
IMU
Attitude
OADC
Sun sensors
Three-axis
control
Reaction Wheels
Torque rods
Solar Panel
Mechanism
Deployment
Dragon Sensors
Structure
System Performance
Requirements
Antennas
Structural
Integrity
Electrical
Housing
Frame Integrity
Multilayer
Insulation
Debris Damage
Outgassing
TPS
Temperature
Control
Passive Control
Degradation
Heat Pipes
Deformation
Optical Solar
Reflectors
Uplink
Antenna
Downlink
Antenna
Oxidation
Communication
GSS
Ground Station
87
Project Management
Statement of Work
The statement of work (SOW) contains the process management must follow during
development. This allows for analysis of the work that needs to be accomplished. It also allows
for accountability. The statement of work begins with a set of requirements that BAI must meet.
SOW Requirements:




BAI shall design a satellite capable of meeting the mission’s objectives and constraints.
BAI shall develop verification methods for all system designs.
BAI shall be responsible for the execution of the verification process.
BAI shall be responsible for the operations of BOSS for the life of the mission.
To ensure that all SOW requirements are met, a work breakdown structure (WBS) must be
created. The WBS supports, manages, and oversees development of BOSS. A WBS has been
developed and is shown in figure 5.
BOSS Project
Project
Management
Budget
Schedule
Deliverables
Ground Segment
Systems
Engineering
System
Requirements
Verification
Ground Support
Operations
Launch
Flight Segment
Subsystem Leads
Design
Integration, Test,
& Verification
Interfaces
Fig 5.
Work breakdown structure. The work breakdown structure shown contains only
the elements which apply to the management of the BOSS project. Subsystem
leads include leads for OADC, TPS, EPS, C&DH, COMM, Payload, and
Structures.
Team Structure
Budget
Launch Services
The Falcon 9 will carry BOSS to its orbit as a secondary payload to minimize costs. At
approximately $4,000 per kilogram, the Falcon 9 offers the best value as a launch vehicle. Figure
2 displays costs comparisons to similar launch vehicles. Once launch services and contingency
are factored into the cost of launch, BOSS will have an estimated launch price of $2.2 million.
89
Figure 2. At a maximum payload, the Falcon 9 offers a substantially lower
launch price than its competitors. 6
Project Cost
A cost analysis was performed for BOSS for the PDR, and has been refined and updated for the
CDR to more accurately represent the actual cost.
Year 1
Year 2
Year 3
Year 4
Year 5
$9,000 / mo
$7,000 / mo
$5,000 / mo
$5,000 / mo
$816,000
$505,000
$250,000
$816,000
$520,000
$250,000
$200,000
$220,000
$200,000
$200,000
$200,000
$200,000
$200,000
30%
30%
30%
30%
$244,800
$151,500
$45,000
$244,800
$156,000
$75,000
$182,400
$126,200
$60,000
$60,000
$60,000
$60,000
$60,000
Salary
Base Rate
Senior Personnel
Junior Personnel
Technician
Admin Personnel
Fringe Benefits
Senior Personnel
Junior Personnel
Technician
Admin Personnel
Materials &
Supplies
Structures
12 mm Screws
16 mm Screws
Raw 6061
Aluminum
$7.95/25
screws
$8.59/25
screws
$1.74/kg
Honeycomb
6061 Aluminum
Approx. $45
per panel
COMM
Transmitter
$12,000
$87.45
$25.77
$65
$65
$270
$12,000
90
Receiver
Antenna
EPS
Solar Arrays
Solar Array
Gimbal
Lithium Ion
Battery
Surrey 150
Power System
Wiring
Payload
Shroud (6)
Sensor and
control box (6)
Hinge (12)
Release
mechanism (5)
Bolt (130)
Washer (130)
Nut (130)
C&DH
On-Board
Computer
COTS Software
TPS
Kapton Film
Aluminized
Teflon Solar
Reflector
Mylar Sheets
Elastisol
Adhesive
Beta Cloth
Netted Spacer
6061 T6
Aluminum Sheet
4mm thick
$10,000
$16,000
$10,000
$16,000
$68,300
$30,000
$68,300
$30,000
$10,650
$10,650
$12,250
$12,250
$10,000
$10,000
$0
$0
$0
$0
~$600
$110
$7200
$550
$12.46 per
pack of 100
$24.92
$5.98 per
pack of 5
$2.77 per
pack of 5
$155.48
$19.39
$64,350
$13,750
$176.18/m2
$11071.90/m2
$338.27
$21258.04
$26.91/m2
$0.34/mL
$51.67
$153.85
$269/m2
$800/coil
$110/
840000mm2
$516
$800
$110.00
Hook and Pile
Tape 25mm wide
OADC
$68.61/
22860mm
$68.61
Accelerometer
Ring Laser
Gyroscope
Sun Sensor (2
$6,545
$15,550
$39,270
$46,650
$102,500
$102,500
91
unit package)
GPS
Reaction Wheels
(4 unit package)
Torque Rods
IMU Computer
De-orbit
module(Nanosat
Terminator
Tape)
6061 T6
Aluminum
(24mm)
6061 T6
Aluminum (6
mm)
M6 #3 Machine
Screws
M3.5 #2
Machine Screws
M5 #2 Machine
Screws
Subcontracts
Payload
Manufacturing
and assembly
OADC
University
Equipment
Structures
Aluminum
manufacturing
equipment (mill,
cnc, etc)
EPS
Misc Electrical
Equipment
Verification/
Testing
COMM
Flight testing
Payload
Vibration testing
Structures
Vibrations Test
Launch Loads
Test
$277,100
$374,500
$277,100
$374,500
$14,000
$34,000
$65,000
$42,000
$34,000
$130,000
$216.00
$432.04
$242.14
$484.28
$5.97
$5.97
$9.66
$9.66
$8.16
$8.16
$40 / hour
$1000
$40,000
$80,000
$3,000
$8,000
$1,000
$2,000
$12,000
$500
$10,000
$40,000
92
General Strength
and Fatigue
testing
C&DH
DITL Testing
Other
Structures
Contingency
Launch Services
Total
Total Cost
$10,000
$500/hr
$2,500
$25,000
$3,396,555
$25,000
$10,000
$2,200,000
$2,171,248 $3,537,000
$10,150 k
$5,000
$5,000
$525,000
$525,000
Mass
Component
COMM
Antenna
Receiver
Transmitter
EPS
Solar Arrays
Solar Array Gimbal
Lithium Ion Battery
Surrey 150 Power System
Wiring
Payload
Shroud (6)
Sensor and control box (6)
Hinge (12)
Nut (130)
Bolt (130)
Washer (130)
Release mechanism (5)
Structures
2 Side Panels (B)
2 Side Panels (A)
2 Top Panels
4 Side Webs
Frame
320 Screws
32 Brackets
Internal Shelf
C&DH
On Board Computer
TPS
Kapton film
Mass (kg)
0.08
1.3
1.0
1.84
1.2
0.786
3.92
1
4.08
13.38
18.00
.6
.8532
.048
1.5
3.376
3.298
4.176
17.92
6.743
1.728
.296
4.51
0.07
0.46
93
Solar Reflectors
Mylar Sheets
Adhesive
Radiation Protection Boxes
OADC
Accelerometer(6 units)
Ring Laser Gyroscope(3 units)
Sun Sensor (2 unit package)
GPS
Reaction Wheels (4 unit package)
Torque Rods(3 units)
IMU Computer
Nanosat Terminator Tape (2 units)
IMU Shroud
IMU Mount
Contingency (20%)
Total
0.149
0.0312
0.581
5
0.426
1.362
0.6
0.95
10.4
3.6
0.7
1.6
2.98
8.29
25.75
154.6
Schedule
To evaluate BAI’s performance, standards must be established. Since project management is
intangible, a verification process may not be possible. Therefore, BOSS’s project management
will be evaluated based on the execution of the established schedule and meeting the budget
constraints
To ensure mission success in a timely matter, a robust schedule containing a critical path to
success will be followed. The schedule contains milestones which represent events such as
launch, deliverables’ due dates, and end of mission. The appendix contains a gant chart which
outlines the schedule for the BOSS project. Table X provides an overview of BOSS’s schedule.
Table X. An overview of the BOSS project’s schedule
Milestone
Deadline
Phase
SIR
09/2014
C
TRR
10/2014
D
ORR
04/2015
D
FRR
11/2015
D
Launch
Q2 2016
E
PLAR
11/2016
E
CERR
06/2017
E
End of Operations
03/2018
DR
06/2018
F
End of Mission
07/2018
-
94
The phases shown in table X represent are defined in NASA’s Systems Engineering Handbook.
However, the BOSS project timeline can be split into segments defined by BAI as follows:
Fabrication: Fabrication lasts approximately nine months. During this time, subsystem
components will be purchased and/or manufactured and assembled. Staff will be employed for
Structures, C&DH, TPS, COMM, OADC, EPS, and Payload to complete this timeline segment;
with most capital being focused on Structures and OADC.
Integration and Testing: Integration and testing will last approximately fourteen months. This
portion of the schedule will incorporate the various subsystems assemblies into the final product.
It also includes verification and testing. Staff will be employed for Structures, C&DH, TPS,
COMM, OADC, EPS, and Payload to complete this timeline segment. The C&DH team will
devote most of their resources towards programming, which will continue to launch.
Operations: Operations begins at launch and lasts approximately 22 months. During this portion
of the project, personnel for GSS and project management will be employed. Operations
concludes at the initiation of de-orbit.
References
1. Kelso, TS, “Analysis of the Iridium 33 and Cosmos 2251 Collision”, Advanced Maui
Optical and Space Surveillance Conference, September, 2009
2. Garcia, Mark, “Space Debris and Human Spacecraft”,
http://www.nasa.gov/mission_pages/station/news/orbital_debris.html#.UnHnuSeMlb4
3. Liou, JC; Adilek, Albert; Burchell, Mark; Corsaro, Robert; Giovane, Frank;
“DRAGONS- A Micrometeoroid and Orbital Debris Impact Sensor” , NASA Orbital
Debris Program Office , 2012
4. Kapurch, Stephen J. “NASA Systems Engineering Handbook”, DIANE Publishing, 2010
5. Tsao, Ngo, Anderson, Sadilek, Pisacane, Giovane, Corsaro, Burchell, Stansberry, Liou;
“Prototyping and Testing a Debris Resistive Acoustic Grid Orbital Navy Sensor” , United
States Naval Academy, 2011
6. Wang, Brian, “Upgraded Spacex Falcon 9.1.1 will launch 25% more than old Falcon 9
and bring price down to $4109 per kilogram to LEO” ,
http://nextbigfuture.com/2013/03/upgraded-spacex-falcon-911-will-launch.html
7. http://orbitaldebris.jsc.nasa.gov/faqs.html
8. Fitz-Coy, N. “Class Notes 8/26/13.” Aerospace Deisgn 1. Fall 2013.
9. Crassidis, John L, & Markley, Flandis (1996, May). Attitude estimation using modified
Rodrigues parameters. In NASA CONFERENCE PUBLICATION (pp. 71-86). NASA.
10. Nagabhushan, Vivek; Fitz-Coy, Norman G. & Johnson, Shawn C. Estimation of
spacecraft angular acceleration using linear accelerometers.
11. IN SITU MEASUREMENT ACTIVITIES AT THE NASA ORBITAL DEBRIS
PROGRAM OFFICE
12. Space Mission Analysis and Design (3rd edition). Microcosm Inc. 1999.
13. http://www.honeybeerobotics.com/flight-programs/44-hinge
95
14. Type 11 biaxial gimbal. (n.d.). Retrieved from
http://www.moog.com/literature/Space_Defense/Spacecraft/Spacecraft_Mechanisms_Pro
duct_Catalog2.pdf
15. Small satellite solar panels datasheet. In (2012). Small Satellite Solar Panels Datasheet.
Clyde Space. Retrieved from http://www.clyde-space.com/documents/2625
16. Power system 150. In (2013). Surrey Satellite Technology LTD. Retrieved from
http://www.sstl.co.uk/getattachment/80c25f64-d725-4907-b942-a3df147cac16/PowerSystem-150
17. Space qualified batteries. In (2011). Clyde Space. Retrieved from http://www.clydespace.com/documents/2343
18. Levins, D. (n.d.). Protection concepts used in spacecraft power systems. Noordwijk,
Netherlands.
19. http://andrews-space.com/torque-rods/
20. F. Beer, R. Johnston, et al., Mechanics of Materials, 6th ed. Boston, MA: McGraw-Hill, 2012,
App. B.
21. http://oce.jpl.nasa.gov/practices/1205.pdf
22. http://www.sncspace.com/ss_space_technologies.php
23. http://spacemath.gsfc.nasa.gov/weekly/3Page25.pdf
96
Appendix
Acronyms
AC – Alternating Current
BAI – Bravo Aerospace Industries
BOSS – Bravo Orbital Space Satellite
C&DH – Command and Data Handling
CMG – Control Moment Gyroscope
COMM – Communications
COTS – Commercial Off-the-Shelf
DC – Direct Current
DOD – Department of Defense
DRAGONS – Debris Resistive Acoustic Grid Orbital Navy Sensor
EPS – Electrical Power System
GEO – Geostationary Orbit
GPS – Global Positioning System
GSS – Ground Support System
IMU – Inertial Measurement Unit
LEO – Low Earth Orbit
MLI – Multi Layer Insulation
OADC – Orbital Attitude Determination and Control
PCDU – Power Conditioning Distribution Unit
PV – Photovoltaic
SOW – Statement of Work
STK – Systems Tool Kit
TPS – Thermal Protection System
TRL – Technology Readiness Level
97
Limited Life Items List
OADC
Reason for
selecting
Item
Life
Requirements
Potential
Impact on
mission
parameters
Item
Part Number
Reaction
Wheel
Surrey 100 SP-0
The satellite
requires threeaxis control.
Torque provided
is sufficient for
this class of
satellites.
At least 2 years
Loss of control
of two or more
wheels will
prevent full
control of
satellite’s
attitude.
Torque Rod
Andrews TQ-40
Momentum bias
that accrues
from the
reaction wheels
must be
dumped.
At least 2 years
Sun Sensor
Surrey 2-Axis
GMC Sun Sensor
Orientation of
satellite relative
to sun will allow
EPS to generate
more power by
orienting the
solar panels.
At least 2 years
Complete
failure of a
torque rod will
allow
momentum bias
to build along
an axis,
preventing
accurate control
of the satellite.
Failure of a sun
sensor will
reduce
knowledge of
position and
may hamper
ability to
position the
solar panels for
maximum
power
generation.
Surrey SGR-20
Knowing
satellite’s
position relative
to GPS satellites
will allow for
determination of
orbital position.
At least 2 years
Accelerometers:
Honeywell
Knowledge of
orientation is
At least 2 years
GPS
Augmented
IMU
Failure of GPS
will reduce onboard
knowledge of
orbital position;
further position
would need to
be determined
by NORAD TLEs.
Failure of any
accelerometer
Plan
Tests of accuracy
of control in
desired
configuration for
conditions more
severe than
expected will be
performed to
determine
efficacy.
Momentum
dumping ability
will be tested on
the ground to
determine efficacy
and optimal
placement on
satellite.
Position data
during all orbits
will be simulated
to make sure that
positions on
satellite are
optimal.
Signal test on
ground to
determine
accuracy and
optimal
placement of
receivers on
satellite.
Test of
configuration to
98
Terminator
Tape
QA3000-010
Gyroscopes:
Honeywell
GG1320
required to
accurately
resolve impacts
on the
DRAGONS
sensor and
orient the
spacecraft; to
advance the
capabilities of
small satellites.
Nanosat
Terminator Tape
Satellite must
deorbit within
25 years of endof-life, this
method will
allow for
deorbiting
within that time
frame.
At least 5 years
will reduce
accuracy of
angular
acceleration
along that axis;
Failure of a
gyroscope will
reduce angular
accuracy.
Complete
failure of IMU
will prevent
accurate
determination
or control of
satellite and will
render it unable
to distinguish
impacts.
Failure to
deploy either
strip will delay
deorbiting of
the satellite,
potentially
beyond the
required
deorbit period.
determine
accuracy
compared to
simulation and to
find optimal
placement inside
the satellite.
Sending deploy
command will be
simulated to
ensure arrival of
signal.
C&DH
Item
Part Number
High
Performance
Flight
Computer
Andrews Model
160
Reason for
selecting
Item
Provided desired
interfaces and
sufficient flash
memory.
Life
Requirements
Potential
Impact on
mission
parameters
Plan
3 years
Provides central
processing for
entire satellite.
DITL testing to
verify system.
Life
Requirements
Potential
Impact on
mission
parameters
Plan
Failure of
transmitter
will result in a
loss of
communication
and data
transmission
We are going
to test the
transmitter at
both high and
low
frequencies
at extreme
COMM
Item
Transmitter
Part Number
Reason for
selecting
Item
The
transmitter
will be used
to send data
to the ground
station at the
desired
At least 2 years
99
frequency
and band-width
Receiver
The receiver will
be used as a way
to stay in
contact with the
ground station.
The receiver will
be used to
receive
emergency
commands as
well as normal
operation
instructions
The patch
antenna is
small and flat
to allow for a
more
aerodynamic
design. It also
allows for a
wider beam
angle
Antenna
At least 2 years
Losing the
receiver will
result in
complete loss
of control with
the satellite.
At least 2 years
The patch
antenna is
the primary
way to send
signals to the
ground
stations
data rates.
We will also
test
transmitting
in hazardous
weather
conditions
We are going to
test the receiver
at both high and
low frequencies at
extreme data
rates. We will also
test
transmitting in
hazardous
weather
conditions
We are going
to test the
antenna at
both high and
low
frequencies
at extreme data
rates.
We will also
test
transmitting
in hazardous
weather
conditions
EPS
Item
Battery
Solar Arrays
Part Number
Reason for
selecting
Item
Battery is
needed to store
power
generated from
the solar arrays
to be used by
the components
of the satellite
when in the
eclipse of the
orbit
Solar arrays are
Life
Requirements
At least 2 years
(Life of Mission)
At least 2 years
Potential
Impact on
mission
parameters
Failure of
battery would
result in
inability to store
power and
therefore
inability to
power
components in
the eclipse of
the orbit
Failure of solar
Plan
Battery storage
test along with
battery
degradation test
along with safe
positioning to
avoid radiation/
interference;
overcurrent
protection
Solar panel
100
used to collect
power from the
sun’s solar rays
to be saved to
the battery to
power the
satellite’s
components
Gimbal will be
used in junction
with attached
hinge to orient
the solar panel
such that the
solar cells are
normal to the
incoming sun
rays
BCM/PDM will
be used to
regulate and
deliver desired
voltages from
battery to
components
Gimbal
Battery
Conditioning
Module/
Power
Distribution
Module
(Life of Mission)
arrays would
result in
inability to
generate power
and prevent all
components
getting power
degradation tests
along with solar
panel stress tests;
radiation,
vibration and
thermal cycling
tests
At least 2 years
(Life of Mission)
Failure of
gimbal would
result in
inability to aim
solar arrays at
the sun and
would severely
limit power
production
Perform
simulations with
loss or partial loss
of power,
vibration and
thermal cycling
tests
At least 2 years
(Life of Mission)
Failure of
BCM/PDM
would result in
incorrect
voltages or no
voltages making
their way to
components
Perform ammeter
test, internal
redundancy, over
voltage protection
system
Payload
Item
DRAGONS
Sensor
Part Number
Reason for
selecting
Item
The sensor will
be used to
collect data on
MMOD and
estimate the size
of the impact,
impact speed,
and impact
direction.
Life
Requirements
At least 2 years
(For the duration
of the mission)
Potential
Impact on
mission
parameters
Failure of a
DRAGONS unit
lessens the
detection area
necessary for
characterizing
the orbital
debris and
limits the
amount of data
being
transmitted for
the mission.
Plan
Because the
sensors are
donated, in-house
tests will be
conducted by the
supplier to ensure
that the sensors
will last the
duration of the
mission.
STK analyses will
be run to ensure
that the estimated
number of hits will
not surpass the
area of the
sensors during the
product’s lifetime.
101
Structures
Item
Part Number
Brackets (32)
Reason for
selecting
Item
Life
Requirements
Component was
needed to hold
the side webs to
the frame of the
satellite during
launch/flight
2 years (Life of
Mission)
Potential
Impact on
mission
parameters
Failure of
brackets could
cause the webs
to detach from
the frame,
resulting in
potential
structural
collapse
Screw failure
could cause the
webs to detach
from the frames
or the panels to
detach from the
rest of the
structure,
leading to
structural
collapse or the
exposure of
internal
components to
the space
environment
See above
Brackets to be
placed under both
load analysis and
also launch
loads/vibrations
testing to ensure
survival during
launch/flight.
Frame
components will
be put through
load analysis
software and also
through launch
loads and
vibrations testing
to ensure
structural integrity
throughout the
mission.
Again, panels will
be extensively
analyzed using
software and
actual
12 mm M6x1.0
Screws
McMaster-Carr256
Used to fasten
the side webs to
the brackets and
frame
2 years (Life of
Mission)
16 mm M6x1.0
Screws
(McMaster-Carr
– 64)
Used to fasten
the panels to the
frame
Used to hold the
satellite
together during
the mission life
2 years (Life of
Mission)
2 years (Life of
Mission)
Frame failure
will lead to the
webs and
panels
collapsing into
satellite, ruining
most electrical
components
and ending our
mission.
Used to shield
the satellite’s
internal
components
from the space
2 years (Life of
Mission)
Panel failure
will expose the
interior of the
satellite to the
space
Frame
Components
(12)
Panels (8)
Plan
Screws are COTS
from McMasterCarr and have
been tested to
withstand launch
loads/vibrations
See above
102
environment.
Honeycombed
6061 Aluminum
used to reduce
mass.
Side Webs (4)
Used to provide
structural
support to the
satellite during
mission as well
as give locations
to attach other
components to
the inside or
outside of
satellite
2 years (Life of
Mission)
environment
and potentially
destroy any
components
attached to that
panel (the
antenna, solar
arrays, payload,
etc.)
Web failure will
decrease
structural
strength at that
side of the
satellite and
could cause the
collapse of its
attached panel.
It could also
lead to the
failure of any
subsystems that
have
components
attached to it
(such as the TPS
sensors or the
sun sensors)
loads/vibrations
tests to ensure
integrity
throughout the
mission life.
Webs will be
placed through
launch loads and
vibrations tests
after passing
computer analysis
to ensure failure
does not occur
during mission.
TPS
Item
Part Number
Multi-Layer
Insulation
Thermistors
Reason for
selecting
Item
The satellite
requires
overall
thermal
protection for
interior
components.
Measurement
Specialties,
Inc. ESCC
Surface
Sensors
Affordable,
lightweight,
small and
space
heritage.
Potential
Life
Impact on
Requirements
mission
parameters
At least 3
Individual
years
components
decrease in
efficiency,
capabilities
or function
when outside
of its
temperature
range.
At least 3
Feedback of
years
component
temperatures
in the event
of
Plan
Test degassing
for durability
and test
thermal cycles
in vacuum
chamber as
well
Momentum
dumping ability
will be tested
on the ground
to determine
103
Series
Radiation
Protection
Shield
4mm thick
840000𝑚𝑚2
6061 T-6
Aluminum
sheet
(McMasterCarr)
Orientation of At least 3
satellite
years
relative to
sun will allow
EPS to
generate
more power
by orienting
the solar
panels.
overheating
they can be
turned off or
change
mode.
Lack of
appropriate
radiation
shielding can
lead to
failure or
malfunction
of electronic
components
efficacy and
optimal
placement on
satellite.
Sheet metal
construction.
Will undergo
vibrational
testing when
coupled with
the satellite.
104
Drawings
EPS
Drawing of the Clyde Space lithium ion battery with dimensions [4]
105
Drawing of the MOOG Type 11 Biaxial Gimbal with dimensions [1]
Drawing of the deployed solar array
106
Drawing of the folded solar array
Payload
107
108
109
110
111
112
Structures
113
114
115
116
117
118
119
Communications
120
OADC
121
122
123
124
Gant Chart
125
126