A Deimos Impact and Observation Spacecraft CubeSat mission with the goal of analyzing the Martian moon Deimos' surface.

1. 1
Having read the Georgia
Institute of Technology
Academic Honor code, I
understand and accept
my responsibility as a
member of the Georgia
Tech Community to
uphold the Academic
Honor Code at all times.
In addition, I understand
my options for reporting
honor violations as
detailed in the code.
Jeff Anderson
Thomas Blachman
Andrew Fallon
John Franklin
Samuel Gaultney
David Habashy
Brian Hardie
Brandon Hing
Zujia Huang
Sung Kim
Jonathan Saenger

2. 2
Executive Summary
A Deimos Impact and Observation Spacecraft (ADIOS) is a microsat with a 6U form factor. The scientific
objective of ADIOS is to analyze the surface and subsurface composition of the martian moon, Deimos. It is
specifically looking for the 1.3 micrometer absorption, and 1.0 - 1.7 micrometer levels for potential prebiotic
volatiles. The former will enable the classification of Deimos as either a Class D or Class C type asteroid, or Mars
ejecta. These are outer asteroids that are thought to have brought prebiotic volatiles to Earth. The latter range is for
finding and determining these prebiotic concentrations to potentially look into how life was created.
ADIOS consists of a six phase mission: launch, hibernation, separation, impact, flyby, and end of life.
Separation will involve the Impactor breaking away from the Observer and continuing on an autonomously guided
course to intercept Deimos. The Observer will pass by one hour after the Impactor hits, to create sufficient time for
the plume to develop. After the plume is analyzed with infrared spectrometry, the data will be relayed to Earth over
the DSN.
The Observer will consist of an infrared spectrometer, chemical propulsion system, transponder, ADCS,
solar panels, and CD&H. The Impactor will consist of GNC, ADCS, inert gas propulsion, and batteries. Both
systems work independently of each other, but the Impactor must hitch a ride with the Observer during launch and
cruise. The GNC on the Impactor consists of a camera guided system. The Impactor occupies 2U while the Observer
is 4U in volume.

3. 3
Table of Contents
Executive Summary Error! Bookmark not defined.
Table ofContents 3
1.0 Mission Objectives and Impact 5
A. Objectives and Specific Success Criteria 5
B. Mission Relevance to Decadal Survey and Impact 5
2.0 Science/Technology Traceability Matrix and Requirements Flowdown 6
3.0 Mission Implementation 7
A. Mission Architecture 7
B. Mission Description 7
C. Trajectory andManeuver Design 7
D. Mission Operations 12
4.0 PayloadSystem 13
5.0 Flight System 17
A. Architecture 17
(1) Structure 17
(2) Electrical Power System 18
(3) Thermal Management 19
(4) Propulsion 20
(5) ADCS 21
(6) GNC 22
(7) Telecommunications 23
(8) C&DH, Flight Software 23
(9) Payload Accommodation 24
6.0 System Engineering 26
A. Technical Resources 26
(1) Mass Budget 26
(2) Power Budget 27
(3) Telecommunications Link Budget 29
(4) Thermal Energy Balance Assessment 30
(5) Data Return Strategy 31
(6) ΔV andPropellant Budget 31
(7) Margins Assessment 33
7.0 Risk Identification & Mitigation 35
8.0 Management,Schedule &Cost 38
A. Management Plan 38

4. 4
B. Program Schedule 39
C. Cost Estimate 42
D. Descope Options 44
Appendix 45
Bibliography 45
Nomenclature 47
Tables and Figures
Description Source Page
Science Traceability Matrix Table I 6
Requirements Flowdown Table II 6
Operational View of ADIOS Mission Figure 1 7
Trajectory Objectives and Tolerances Table III 8
Optimal Trajectory Sketch Figure 2 9
Mars 2020 Position and Velocity Vectors Table IV 9
GMAT View of Flyby Figure 3 10
Trajectory Statistics for Example Separation Dates Table V 11
∆V Contributions by Separation Date Figure 4 11
Change of Required insertion ∆V Figure 5 11
Scenarios for Collision with Deimos Figure 6 12
ADIOS Flight Schedule Table VI 12
Absorption Strengths of Volatiles Figure 7 13
Absorption Strengths of C and D Class Asteroids Figure 8 13
Mass Spectrometer Trade Study Table VII 15
Impactor Architecture Trade Study Table VIII 15
Impactor Power Trade Study Table IX 16
Plume Analysis: Mass vs. Impact Velocity Figure 7 16
Plume Analysis: Debris Altitude vs. Impact Velocity Figure 8 16
Observer and Impactor Modules Figure 9 17
Various Configurations of ADIOS Figure 10 17
Module Frames Figure 11 18
Power Trade Study Table X 19
Thermal Analysis Table XI 19
Temperature Sensitive Components Table XII 20
Propulsion Architecture Analysis Table XIII 20
Release Mechanism Location Figure 12 24
Release Mechanism Figure 13 24
Observer Mass Budget Table XIV 25
Impactor Mass Budget Table XV 26
Spacecraft Mass Budget Table XVI 26
Observer Power Budget Table XVII 27
Observer Power Budget Timeline and Modes Figure 14 27
Impactor Power Budget Table XVIII 28
Downlink Budget Table XIX 28
Observer Operational Temperatures Table XX 29
Impactor Operational Temperatures Table XXI 29
∆V Requirements by Launch Date Figure 15 30
Variation of ∆V Requirements Table XXII 31
Impact Velocity vs. Total ∆V Required Figure 16 31

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Margins Assessment Table XXIII 32
Mission Risk Matrix Figure 17 33
Project-Level Risk Matrix Figure 18 34
Work Breakdown Structure Figure 19 36
Program Schedule Figure 20 37
Critical Path Figure 21 39
Mission Cost Budget Figure 22 40
Cost Breakdown Figure 23 42

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1.0 Mission Objectives and Impact
A. Objectives and Specific Success Criteria
Objectives:
● The Impactor shall collide with Deimos’ surface and generate a plume sufficient enough in size for the
CubeSat Spectrometer to detect.
● The Impactor shall release from the Observer and penetrate Deimos’ surface deep enough to expose
subsurface volatile compounds including oxygen, carbon dioxide, carbon monoxide, water, and ammonia.
● The CubeSat shall analyze the plume with a spectrometer and determine the 1.3 µm absorption levels, as
well as the absorption levels of volatiles and successfully relay this data back to Earth.
Specific Success Criteria:
● Analyze the plume with a spectrometer and determine the 1.3 µm absorption levels, as well as the
absorption levels of volatiles and successfully relay this data back to Earth.
Partial Success Criteria:
● Analyze the surface of Deimos with a spectrometer and determine the 1.3 µm absorption levels, as well as
the absorption levels of volatiles and successfully relay this data back to Earth.
● Or, the Impactor shall release from the Observer and penetrate Deimos’ surface deep enough to expose
subsurface volatile compounds including oxygen, carbon dioxide, carbon monoxide, water, and ammonia.
B. Mission Relevance to Decadal Survey and Impact
Within NASA’s Decadal Survey there is a strong call for a mission to Deimos. Below is an excerpt from the survey.
“Resolving the debate concerning the compositions (and likely origins) of the martian moons Phobos and Deimos
may be relevant to understanding the early history of Mars...Investigation of Phobos and Deimos crosscuts
disciplines of planetary science including the nature of primitive asteroids, formation of the terrestrial planets, and
astrobiology. Key science questions are the moons’ compositions, origins, and relationship to other solar system
materials.”
Due to this interest, Deimos was chosen as the primary focus of the mission proposal. Two major questions arise
from the survey that our mission intends to answer. The first, from the decadal survey, “Are the moons possibly re-
accreted Mars ejecta? Or are they possibly related to primitive, D-type bodies?” For this reason our mission seeks
to determine whether Deimos shows strong absorption in the 1.3 µm, which would denote Mars ejecta, or low which
would indicate a D-class asteroid. The second question,“What does [D-Type asteroids] concentration oforganic
volatilessuch, methane, water, ammonia and carbon dioxide,tell us about how organics were distributed in the
early solar system.” Our mission seeks to determine theses volatiles by analysing the 1.0 - 1.7 µm. To link to the
decadal survey it calls to, “Characterize organic chemistry, including (where possible) stable isotopic composition
and stereochemical configuration.Characterize co-occurring concentrations ofpossible bioessential elements.”
Finally, the decadal survey directly calls for a mission that matches our science objectives, “These questions can be
investigated by a Discovery-class mission that includes measurements of bulk properties and internal structure,
high-resolution imaging ofsurface morphology and spectral properties, and measurements of elemental and
mineral composition.”

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2.0 Science/Technology Traceability Matrix and Requirements
Flowdown
Mission Objective: Measure the internal subsurface composition ofDeimos to determine its origins
and organic volatile levels
Table I. Science Traceability Matrix
Science
Objectives
Measurement
Objectives
Measurement
Requirements
Instrument
Requirements
Instruments Data Products
Deimos
Internal
Composition
Measure ratio of
iron in internal
composition
Spectronomy for
160 seconds
Measure the 1.3 µm
absorption levels of the
plume
ARGUS
Spectrometer
Graphs of
Spectronomy
Readings
Internal Volatiles
Determine the
amount and type of
subsurface volatiles
Spectronomy for
160 seconds
Measure the 1.0 µm -
1.63 µm. absorption
levels of the plume
ARGUS
Spectrometer
Graphs of
Spectronomy
Readings
Table II. Requirements Flowdown
Determine
Subsurface
Composition
of Deimos
Impactor
Shall
Navigate to
Deimos
- GNC must determine spacecraft position
- ADCS must keep propulsion pointed in
correct direction
- Propulsion must have fuel for necessary
aligned maneuvers
- Impactor position knowledge < 3 km
- Impactor pointing knowledge < 0.1°
- Impactor pointing accuracy < 0.1°
- Impactor ∆V > 35 m/s
Observer
Shall Fly by
Deimos
Plume After
Impact
- GNC must determine spacecraft position
- ADCS must keep propulsion pointed in
correct direction
- Propulsion must have fuel for necessary
aligned maneuvers
- Observer position knowledge < 600 km
- Observer pointing knowledge < 0.15°
- Observer pointing accuracy < 0.15°
- Observer ∆V > 218 m/s
Observer
Shall Analyze
Plume
Composition
- ADCS must keep spectrometer pointed
at plume
- Spectrometer must analyze plume
- Pointing knowledge < 0.15°
- Pointing accuracy < 0.15°
- Spectrometer effective range < 600km
- Spectrometer wavelength range = 1.0 to 1.7
μm
Observer
Shall
Transmit
Back Data
- Spacecraft must communicate with DSN
on Earth
- Eb/No > 6

8. 8
3.0 Mission Implementation
A. Mission Architecture
Figure 1. Operational View of ADIOS Mission
Following the separation from Mars 2020, the mission will be composed of five key periods. First the
spacecraft will perform any necessary alignments and hibernate. Once the spacecraft enters the Mars sphere of
influence, the Observer and Impactor will separate. The Observer will slow down and the Impactor will continue on
its collision course to Deimos. The Impactor will hit Deimos and the Observer will fly by Deimos approximately
one hour later to analyze the plume using spectrometry. The final action is for the spacecraft to transmit the data
back to earth.
B. Mission Description
The ADIOS spacecraft will rideshare on the Mars 2020 mission to a Mars trajectory. The spacecraft will be
composed of an Impactor module and an Observer module. The Impactor module will collide with Deimos,
generating a plume. The Observer will perform a flyby of Deimos when the plume has reached maximum height and
will perform spectrometry on the plume. The spectrometry data gathered will then be transmitted back to earth via
the DSN.
C. Trajectory and Maneuver Design
(1) Key Requirements
There were several things to consider when designing the trajectory. In order to be considered a success,
the trajectory must meet the following requirements:
1. The overall ΔV budget of mission shall not exceed the capabilities of our propulsion system.
2. The Earth-Mars transfer orbit shall reach Mars at a minimal inclination and a reasonable flyby offset in order
to minimize ΔV required to impact Deimos.
3. The Impactor’s final trajectory shall cross into Deimos with an error less than the Moon’s diameter.
4. The Impactor shall impact Deimos at a speed no less than 3.5 km/s to produce a sufficient plume size.
5. The Observer shall pass Deimos at a safe minimum altitude, at least 3 minutes later than impact.
6. The Observer shall not crash into Deimos, Mars, or any other body before transmitting its data.

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(2) Transfer and Maneuver Strategy
Provided a nominal approach trajectory from Mars 2020 before detachment, the first
burn will take place just four days after separation. This maximizes the effectiveness of
the first maneuver to fine tune the desired altitude and inclination. This burn also
matches time of arrival at Mars’ SOI to a desired value. The second burn will occur after
the long cruise phase en route to Mars at its SOI (approximately 600,000 km altitude).
This maneuver will place the entire CubeSat on a collision course with Deimos.
Immediately afterwards, the impactor section of the CubeSat will separate and will
continue towards Deimos. The observer will then execute another burn mainly to reduce
its velocity. This is because the observer must arrive at Deimos one hour after impact in
order for the plume to fully develop. Once the observer reaches data acquisition range, it
will need at least 160 seconds to record the data. Getting within this range within this
timeframe is the key to mission success. The command tree for the trajectory solver and
values can be seen to the left.
Table III. Achieve goals with values and tolerances
Goal Value Tolerance
Achieve 𝐵 ∘ 𝐵 -30,000 km 0.01
Achieve 𝐵 ∘ 𝐵 0 km 0.01
Achieve SOI 600,000 km 0.1
Achieve Impact 3 km 0.001
Achieve Position 1 -100 km 0.001
Achieve Position 2 100 km 0.001
Achieve Position 3 0 km 0.001
The purpose of the value of ‘𝐵 ∘ 𝐵’ is to position the CubeSat between Mars
and the Sun. It indicates the offset of the flyby trajectory the spacecraft follows upon
entering Mars’ SOI. To minimize impact burn ΔVi required to impact Deimos and
maximize impact velocity, the most desirable trajectory is a Mars-flyby trajectory tangent to Deimos’s orbit at its
periapsis. To estimate the required offset, it is first estimated that the upon arrival at Mars of SOI, spacecraft speed is
approximately 2.67 km/s. This was found using the specific energy equation:
(1)
With the altitude (r) equal to 600,000 km, velocity (v) equal to 2.67 km/s, and the semimajor axis (a) equal to -
6,148 km. For hyperbolic orbit, the periapsis radius is given as:
(2)
Where b is the initial offset. Equating rp with Deimos’ semi-major axis (23,463 km), the offset is found to be
approximately 30,000 km. Since a retrograde orbit is desired, 𝐵 ∘ 𝐵 is thus -30,000 km. Figure 2 shows a sketch of
optimal offset and positioning.

10. 10
Figure 2. Optimal Mars-SOI entrance offset and impact trajectory
On the other hand, ‘𝐵 ∘ 𝐵’ is set to 0 to give an equatorial trajectory, since the inclination of Deimos’s orbit
relative to Mars is small. The position values get the Observer within a specific distance of Deimos so the data
acquisition range can be guaranteed. In Mars Approach Propagation, “Prop to Specific Date” also allows adjustment
of time-of-arrival at Mars SOI, the effects of which are discussed in (3) ΔV Budget and Optimization.
The GMAT simulation starts at when the spacecraft separates from Mars 2020 in Earth-Mars transfer orbit. A
arbitrary value of such condition is given in Table IV, since we have no control over the exact trajectory of Mars
2020 and we require the Earth-escape burn. This initial condition will ensure the spacecraft arrives near Mars in
reasonable time, but mid-transfer maneuvers will likely be needed.
Table IV. Approximate Initial Mars 2020 Trajectory, Earth MJ2000, Equatorial, Cartesian, Epoch: 15 Jul 2020
12:00:00
x (km) 95866
y (km) 21709
z (km) 25756
Vx (km/s) 3.019
Vy (km/s) 0.822
Vz (km/s) 2.313

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Figure 3. View from Deimos of Observer flyby, Red and green is Observer, Blue is Impactor.
(3) ΔV Budget and Optimization
For a given condition upon arrival of Mars’s SOI, Deimos must be at the optimal position to minimize the
ΔV required for the spacecraft to impact Deimos, i.e. it must be crossing the spacecraft’s periapsis with respect to
Mars at the time the spacecraft reaches periapsis. However, this is not always achievable since the launch of Mars
2020 mission does not consider any contact with Deimos. Given that Deimos’ orbital period around Mars is
approximately 30.32 hours, we must make sure that:
1. The spacecraft is able to delay, or shift earlier, the time of arrival by at least 15 hours.
2. The spacecraft has an acceptable window within each period to impact with Deimos.
Since we do not have control over the exact launch time, we could only set an arbitrary starting condition
and adapt maneuvers to it. For any given starting time, the TCM will ensure the arrival time is on February 17th,
2021 as an example impact date. Table V shows some example schedules of the flight path. Figure 4 shows the total
relationship between the amount of ΔV required and separation date. Each burn’s contribution to the total ΔVis also
shown.
On the other hand, impact of arrival time on required ΔVc is significant. This can be seen in Figure 5,
showing variation of minimum required impactor maneuver to impact Deimos and corresponding impact velocity,
given a certain initial velocity and position upon entrance of Mars’ SOI but at a different time. For example, for a
departure (initial separation) time of 7/23/2020 12:00:00, the ideal time of arrival at Mars SOI is around 2/14/2021
17:00:00. The corresponding ΔVc is only 6 m/s for this condition and the impact velocity is 4.61 km/s. There are
about 12 hours (40%) over each period when the required ΔVc is less than 50 m/s, and over 20 hours (67%) each
period when the achievable impact velocity is greater than 4 km/s. Some impact trajectories corresponding to
different time-of-arrival at Mars SOI are shown in Figure 6, green trajectories representing less than 20 m/s of ΔVc
required.
Therefore, although the on-board algorithm should aim for the optimal time of arrival, there will still be a
good tolerance of error. It should be noted that if the time of arrival falls out of these windows, the fuel for Observer
slow-down burn (ΔVo) will be compromised; in some extreme cases (less than 15% time window), the spacecraft
may not have enough fuel to even impact Deimos.

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If the spacecraft successfully matches the time of arrival with optimal target, ΔVo will comprise the most
part of the total ΔV budget. However, ΔVo is very flexible with the lead time of Impactor-Observer separation and
desired time-of-arrival difference (at Deimos). As shown in Table V, the ΔVT required will be generally less than
200 m/s and can be as low as 135 m/s.
Table V. Trajectory Statistics for Example Separation Dates
Separation
Date
Impact
Date
Flyby Date
ΔVi
(m/s)
ΔVc
(m/s)
ΔVo
(m/s)
ΔVT
(m/s)
Impact V
(km/s)
Fuel Used
(kg)
7/21/2020
12:00
2/17/2021
7:18
2/17/2021 8:18 55.125 16.172 71.235 142.532 4.592 0.676
7/22/2020
12:00
2/17/2021
7:32
2/17/2021 8:32 53.718 13.700 69.804 137.222 4.587 0.657
7/23/2020
12:00
2/17/2021
7:44
2/17/2021 8:44 51.821 11.835 68.773 132.429 4.581 0.638
7/24/2020
12:00
2/17/2021
7:55
2/17/2021 8:55 49.487 10.327 68.049 127.863 4.574 0.621
Figure 4. ΔV Contributions by Separation Date
Figure 5. Change of required insertion ΔVc over approximately one Deimos orbital period

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Figure 6. A few cases of the collision course over one 30-hour period for impact
D. Mission Operations
Once ADIOS separates from Mars 2020, it will hibernate for approximately four days before conducting its
first burn, ΔVi in order to obtain a 𝐵 ∘ 𝐵 of -30,000 km as described in detail earlier. Several small burns will take
place to maintain this trajectory. After about 200 days, or until ADIOS reaches Mars’ SOI, the second burn, ΔVi,
will take place forcing ADIOS into a collision course with Deimos. Immediately after the burn, ADIOS will separate
into an ‘Observer’ and ‘Impactor’ craft. The Impactor will continue on its current trajectory. The Observer instead
performs another burn, ΔVo, to delay its flyby of Deimos by about an hour so that the plume reaches maximumsize.
During the flyby, at least 160 seconds is required for the spectrometer to complete its data acquisition.
Table VI depicts an example flight schedule for several separation dates. The ‘Observe Time’ column
depicts the amount of time in seconds that the Observer will be within 600 km of Deimos, which is the maximum
range of the spectrometer.
Table VI. Flight Schedule
Days Elapsed since Separation
Separation
Date
ΔVi ΔVc + ΔVo Impact Flyby
Observe Time
(seconds)
21-Jul 2020 4 208.22667 210.80 210.85 248.72
22-Jul 2020 4 207.22667 209.81 209.86 260.54
23-Jul 2020 4 206.22667 208.82 208.86 257.12
24-Jul 2020 4 205.22667 207.83 207.87 255.70

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4.0 Payload System
The payload of the ADIOS mission consists of a few different instruments for both the Observer and the Impactor.
Key Requirements for Instrument Selection
Spectrometer
The spectrometer must be able to detect the subsurface volatiles. A lack of these volatiles would indicate a
hot birth of Deimos, and would point to it being ejecta from Mars. The subsurface volatiles are made of chemicals
such as: Water, CH4 ,CO2 CO. In order to accomplish this, the spectrometer must be able to make measurements
within the range of 1.0 and 1.7 µm, as seen below in Figure 7.
Figure 7. Absorption strengths of volatiles (Argus)
The spectrometer analysis must be able to determine if Deimos is a captured asteroid. Objects from the
asteroid belt are primarily categorized as C and D class. These asteroids have very weak absorption spectra in the
1.3µm range, as shown in Figure 6 below.
Figure 8. Absorption strengths of C and D class asteroids (Cloutis)
The spectrometer is required to operate in these ranges if it is to determine the origin of Deimos.

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Impactor
The Impactor needs to be able to create not only a plume big enough to be seen by the spectrometer, but
also eject material deep enough to expose any subsurface volatiles. Additionally, the pointing accuracy of the
cubesat is 0.003° and the field-of-view (FOV) of the Argus spectrometer is 0.15°. This means that from the
instrument’s 600 km maximum range the plume must be greater than 0.78 km wide to guarantee that the plume will
be detected.
(3)
(4)
(5)
(6)
The equations above are the non-dimensional equations used to estimate the yield of the Deep Impact mission from
small scale models (Richardson). From these equations we can determine the size of the crater and therefore the
amount of material ejected. Using the values for the Impactor speed and mass, 4.5 km/s and 4 kg respectively, it can
be found that the plume’s width is 4 km and that material from as deep as 1.57 meters is ejected. These results
produce a margin over the specified requirements, however these values have already been budgeted and the extra
capability will be kept as a descope option.
The velocity of the plume will be fairly slow. The equation below was used to determine the velocity
distribution of the plume as a function of location in the crater.
𝐵 = √𝐵𝐵 ∗ 0.62(
𝐵
𝐵
) −0.255 + 0.06
(7)
Using this equation it was found that only 5% of the ejecta moves faster than the escape velocity of 5.56 m/s. The
particles with the highest altitude will began to fall back to Deimos 50 minutes after impact, marking the maximum
size of the plume. Because of this, a driving requirement is that the Impactor must be able to reach Deimos 50
minutes in advance of the Observer.
Spectrometer Trade Study
The spectrometers that we evaluated were the ARGUS, mini-INMS, and BIRCHES spectrometers. For the
selection of the spectrometer, a trade study was done in order to choose between these three different spectrometers
within the allocated 4U size constraint. The criteria between the spectrometers included mass, volume, power,
spectrum ranges, operating range, data collection rate, pointing, and cost. Table VII below shows the comparison of
the spectrometers with the selected criteria. The cells that are highlighted in green are the most desirable values for
the criteria. With the smallest mass, volume, and lowest power requirements, the ARGUS spectrometer provides the
greatest flexibility and size efficiency for the mission. In addition to these advantages, this spectrometer has a high
TRL. Therefore, ARGUS will be chosen as the main science instrument due to its overall benefit to the mission,
especially since it can detect spectral bands in the 1000 to 1700 nanometer range.

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Table VII. Mass Spectrometer Trade Study
Instrument ARGUS mini-INMS BIRCHES
Mass (g) 230 600 2000
Volume (U) 0.18 1.1 1.5
Power (W) 1.4 1.8 5
Spectrum (nm) 900-2500 0-2000 0-4000
Range (km) 600 675.924 100
Data Rate (Mb/s) 1 0.0013 10
FOV (degrees) 0.15 20 12
Cost ($) 49,500 15,000 200,000
TRL 8/9 6/7 8/9
Impactor Trade Study
The Impactor was already evaluated in a trade study in our mission planning phase. However some options
related to its operation needed to be evaluated. The first was the amount of time spent in its separation period where
it gains its 50 minute gap between itself and the Observer. As seen in Table VIII below the earlier it detaches the
less fuel it needs for the separation burn, but the more power will have to be supplied. The risk also increases the
more time it is required to navigate by itself. The decision was made to release it earlier despite the risk, as it limits
the wet mass of the Impactor that isn’t transferred to Deimos.
Table VIII. Impactor Architecture Trade Study
Time Released Before
Impact (hours)
ΔV Required (m/s)
Weight of Fuel
(kg)
Approximate Power Required
(Wh)
Risk (1-10)
24 130 1.2 60 7
48 65 0.7 120 6
72 30 0.43 180 6
A second decision we made regarded the power source of the impactor. Solar was considered as it would provide
power indefinitely. However, it was found to be much larger and heavier than batteries. So large that it would
require folding solar panels that would raise risk. Batteries were chosen due to their density, and because the
Impactor will only need power for 72 hours no matter what.
Table IX. Impactor Power Trade Study
Power Source Size (cm) Weight (kg) TRL
Solar 38.1 x 31.5 x 10.2 2.359 8
Battery 7 x 3 x 2 - 7 x 3 x 6 0.032 - 0.096 9

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Figure 7. Plume Analysis: Mass ejected as a function of impact velocity
Figure 8. Plume Analysis: Debris altitude as a function of impact velocity
Figures 7 and 8 above are breakdowns of plume size as velocity and material properties are changed. The variables
Cd and b are non dimensional quantities that represent material strength. In the plots the b and cd are varied fromthe
values that represent loose sand (b = 0.165, Cd = 1.54) and solid rock (b = 0.2134, Cd = 2.54) represented as
different colored lines. At the levels of loose sand both the size of the plume and the amount of material ejected are
lowered, with the highest amount being at the values of solid rock. The red dot on the chart represents the values at
which we predict out plume will be generated. As can be seen in the figures the predictions are at the lower point in
the curve, meaning that if there is error in the predictions the result will probably be higher.

18. 18
5.0 Flight System
A. Architecture
The 6U spacecraft is divided into two self-contained and self-operating modules. The larger of the two is the 4U
Observer module. This module contains ADCS, C&DH, comms, EPS, propulsion, and our payload of the
spectrometer. The 2U Impactor module contains all the same subsystems, but instead of the spectrometer there is a
camera which is used for GNC. The two modules are joined by a separation mechanism, which is controlled by the
Observer module. Figure 9 shows the spacecraft with the two modules slightly separated. Note that the radiation
shielding and solar panels are not pictured. Figure 10 shows three different configurations of the spacecraft: with
solar panels stowed, with solar panels deployed, and with the Impactor module separating.
Figure 9. Observer and Impactor Modules, separated
Figure 10. Various configurations of spacecraft
(1) Structure
The structure of the spacecraft is custom-built aluminum. Using an off-the-shelf frame is not an option
because of the size of the ADCS, IRIS V2 comms system, and propulsion system. The Observer frame is composed

19. 19
of 1/2U aluminum frames attached to the central propulsion frame. The Impactor frame is a single aluminum piece.
For both the Observer and Impactor the individual components slot into their respective frames. For example, the
batteries and EPS control boards slide directly directly into the Impactor frame and are bolted in. Figure 11 shows
the Impactor and Observer frames with and without the other spacecraft components.
Figure 11. Module frames, with and without internal components
(2) Electrical Power System
Observer
After conducting a trade study on the power source options shown in Table X, the Observer will rely on
Solar panels to generate power and a 60Wh battery. It will use two Clyde Space Double-Deployable, Double-Sided
Solar Panels. These will line the outside of the cubesat and fit within our structure. This will provide 40 W of peak
power at Deimos (assuming 48% solar irradiance of Earth) and 20.8 W (EOL) of Orbital Power. These provide the
best combination of size and power output for our requirements. The power system will be managed by a Clyde
Space FlexU Cubesat EPS, which is up to 98% efficient regulating 3.3V and 5V Busses. This will handle all the
power conversion requirements for our Observer. The Battery will be a Clyde Space 60Wh standalone battery with
custom battery protection circuitry. This battery option provides better discharge control than regular lithium ion
battery options.
Impactor
For the Impactor, three Clyde Space 40Wh batteries will be used over solar power. Again, after conducting
the trade study, the Impactor would require a lot of power in a short amount of time but is limited on volume, thus
the short time that it needs to be used couple with the small size of the Impactor makes solar infeasible. These
batteries will provide the Impactor with 120Wh, which meets the power requirements. The batteries will be
accompanied by another FlexU Cubesat EPS.
Table X. Trade Study for Power Sources for Observer and Impactor
Observer
Double Deployable Solar
Panels (with Batteries)
Batteries Solar Array Deployment
Lifetime (Years) 10 3 10

20. 20
Mass (kg) 0.8 0.25 1.1
Volume (cm x cm x cm) 200 x 48 x 5 95 x 90 x 27.4 100 x 45 x 100
Peak Power (W) 40 80 90
Total Power (Wh) - 80 -
Impactor
Solar Panels Batteries
Lifetime (Years) 10 3
Mass (kg) 0.8 0.25
Volume (cm x cm x cm) 200 x 48 x 5 95 x 90 x 27.4
Total Power (Wh) - 80
Power Production Delay Yes No
(3) Thermal Management
Equations 7 and 8 below are used to balance the thermal energy of the spacecraft. The top equation Qe is
the energy radiated by the spacecraft. The bottomequation Qa is energy absorbed by the spacecraft plus the energy
added by power used in the spacecraft components. By setting the equations equal to one another and solving for Tr
the equilibrium temperature of the spacecraft can be found.
𝐵 𝐵 = 𝐵 ⋅ 𝐵 ⋅ 𝐵 𝐵 ⋅ 𝐵 𝐵
4
(8)
𝐵 𝐵 = 𝐵 𝐵 ⋅ 𝐵 ⋅ 𝐵 𝐵𝐵𝐵(𝐵) + 𝐵𝐵𝐵𝐵𝐵 + 𝐵𝐵𝐵𝐵𝐵𝐵 (9)
In Table XI below, the maximum and minimum temperatures at key times during the mission are calculated and the
variables used are included. The calculations are also done for the Observer and Impactor together and apart. This is
done because the ratio of sunface surface area and radiation area changes.
Table XI. Thermal Analysis
Observer + Impactor Observer Impactor
α = Absorbed 0.92 0.92 0.92
ε = Emitted 0.85 0.85 0.79
So = Earth Solar Flux 1370 1370 1370
So = Mars Solar Flux 608.9 608.9 608.9
A = Area Absorbed 0.06 0.04 0.04
Ar = Area Emitted 0.22 0.2 0.1
σ = Constant 5.67E-8 5.67E-8 5.67E-8
Watts (min) 25.69 25.69 0.55
Watts (max) 26 52 14.55
Watts (heater) 0 10 0
Earth Cruise (C°) 37.652
Mars Cruise (C°) 0.4701 11.499 -8.67
Mars Full Power (C°) 0.8268 23.3843 16.4572
Below in Table XII is the maximum and minimum tolerable temperatures for the Observer and Impactor
and both together. It lists the components that will first experience negative effects and therefore set the temperature
limits. Without thermal control the spacecraft is able to maintain operable temperatures for all time of the mission
except for when the Observer first separated from the Impactor is is still in cruise. For this part of the flight a small
heater will be used.

21. 21
Table XII. Temperature Sensitive Components
Configuration
Maximum Tolerable
Temperature (°C)
Component
Minimum Tolerable
Temperature (°C)
Component
Observer + Impactor 40 Argus Spectrometer 5 Rocketdyne MPS-130
Observer 40 Argus Spectrometer 5 Rocketdyne MPS-130
Impactor 40 Clyde Space Battery -10
VACCO End-
Mounted MiPS
For thermal management the 10 watts of heat will be added by a small 10 watt thermal kapton strip placed
on the Observer's Rocketdyne MPS-130 to keep it warm enough for operation. The spacecraft stays cool enough
thought the mission, so only passively radiative effects will be used for heat.
(4) Propulsion
The ADIOS mission requires ∆V maneuvers that have no precedence for
CubeSat missions and therefore a reliance on developing technologies is required
and will be discussed in the following paragraphs. Additionally, commercial off-
the-shelf components were favored over custom systems due to cost and time
constraints.
Observer
The Observer module will need to achieve a total ∆V of 218 m/s, in the worst-
case scenario. The fly-by nature of the mission confines the orbital maneuvers to
be carried out in relatively short time periods. This constraint suggests the
elimination of electric propulsion and a solar sail as feasible propulsive methods
after entering Mars’ SOI, as they require long time spans to achieve high ∆Vs. Inert
gas propulsion has the advantage of simplicity but with an average Isp of 70 seconds,
will not produce a sufficient ∆V for the Observer’s maneuvers.26 A chemical propulsion
system provides the greatest confidence of achieving the mission’s objectives as they are largely dependent on the
spacecraft’s ability to adjust its trajectory when approaching Deimos. However, the idea of using a combination of
electric and chemical propulsion for the Observer was explored. An analysis was carried out in GMAT and the
results are displayed in Table XIII. A chemical propulsion system was compared with a combination of chemical
and ion thrusters. The combined architecture envisioned a scenario where the electric propulsion would provide the
initial trajectory and the chemical propulsion system would be used once within Mars’ SOI. These calculations
assumed a separation date from the Mars 2020 spacecraft of July 23, 2020. Additionally, specific impulses of 800
seconds and 240 seconds were used for the ion thruster and chemical propellant system, respectively. As shown
below, the combined system only reduces the necessary fuel by 65 grams. This difference is not enough to warrant
utilizing this option, as two different propulsion systems will only increase power requirements, cost, complexity
and risk. This leaves chemical propulsion as the best option for the mission.
Table XIII. Propulsion Architecture Analysis
Architecture ∆V­­i (m/s) ∆V­­c (m/s) ∆V­­o (m/s) Ion Fuel (kg)
Chemical Fuel
(kg)
Ion and Chemical 43.66 18.25 72.42 0.095 0.47
Chemical 43.68 18.26 72.41 0 0.63
Bipropellant chemical systems are too mechanically complex to scale down to a small satellite systemand
therefore a monopropellant systemwill be used.The propellant that will be used is a hydroxyl ammonium nitrate
fuel and oxidizer blend, commonly known as AF-M315E. It was selected over hydrazine for several reasons.Most

22. 22
importantly, AF-M315E has a significantly lower toxicity level when compared to hydrazine.26 This will reduce
range operations costs,as loading procedures can occur more rapidly and will eliminate the need of the waiver that
is normally required with hazardous,secondary payloads.These loading operations will cost a fraction of the
average contractualcost to load conventionalpropellants, which is $135,000 per NASA mission.26 A lower toxicity
level eliminates the need for redundant components as a propellant leakage is rated as critical rather than
catastrophic,as it is with hydrazine leaks.26 This will yield simpler architectures and provide further cost savings.
Additionally, AF-M315E is 45 percent denser than hydrazine which allows for smaller propellant tanks.26 While
precautions must be taken to ensure hydrazine does not freeze during long periods of coasting, AF-M315E has a
glass transition and cannot freeze.26 This reduces the complexity and cost the thermal management subsystem.This
propellant has an expected Isp of 240 seconds,which is a 12 percent increase over hydrazine.26
Aerojet Rocketdyne’s Modular Propulsion System (MPS-130) will be used as the main system. It will
occupy 2U (10 cm x 10 cm x 22.4 cm) and is equipped with four 1.5 N thrusters.44 The wet mass will total 3.5 kg
and the amount of propellant it will be able to hold is 1.3 kg.44 Operation pressure ranges from 34.5 to 5.9 bar and
operational temperatures range from 5 to 50 Celsius.44 The systemrequires 7 watts to operate the catalyst bed heater
and 4 watts for the thrustervalve startup,while operational power will amount to 1 watt.11 Assuming an initial total
mass of 13.777 kg and using the ideal rocket equation shown in Equation 9, the available
∆V is 233.3 m/s for this configuration.
𝛥 𝐵 = 𝐵 𝐵𝐵 𝐵0 𝐵𝐵(
𝐵0
𝐵 𝐵
) (10)
This will allow for a 7.01 percent margin of ∆V in the event of emergency maneuvers. This
system has a TRL 6 and will be flight tested in early 2017 when NASA’s GPIM
launches.44,45 This mission is intended to demonstrate the viability of using AF-M315E as an effective propellant
for orbital maneuvers. Its successfulcompletion will result in TRL 7+ for the propellant system.
Impactor
While the Impactor will separate from the Observer on course for impact
with Deimos, we have included an inert gas propulsion system for any slight
corrections during its final trajectory. Unknown perturbations cannot be
accounted for in our GMAT analysis, however, the Deep Impact mission allocated a
∆V of 25 m/s for its impactor’s 24-hour flight. Therefore, we expect no more than 35
m/s of ∆V to be required for the ADIOS Impactor.59 Chemical propulsion could introduce the
possibility of contaminating the plume samples and would occupy most of the Impactor’s volume. Electric
propulsion would not allow for quick corrections in attitude and trajectory. Therefore, we have chosen VACCO’s
End-Mounted Standard Micro-Propulsion System (MiPS). This TRL 6 component occupies 0.5U and is equipped
with five thrusters that provide a total impulse of 166 N-s.55 The non-toxic R314a propellant has an Isp of 40
seconds.55 The propellant will self-pressurize over the normal operating temperatures of 0 to 60 Celsius.55 The
component houses a proportional heater and temperature sensorfor adjusts in propellant pressure.55 The maximum
standby and steady-state powerconsumptions are 0.25 watts and 5 watts, respectively.55 Wet mass for the systemis
0.924 kg and it will hold 0.423 kg of propellant.55 Using Equation 9 with these specifications and assuming a total
Impactor mass of 4.381 kg, this system is expected to supply a ∆V of 39.8 m/s. This provides a
13.85 percent margin for additional maneuvers.
(5) ADCS
The Attitude Determination and Control System (ADCS) provides three-axis
attitude control for all phases of the mission to meet pointing and instrument
needs. The method of ADC control for the Impactor and Observer will be through
the Blue Canyon Technologies (BCT) XACT Attitude Control System. The BCT XACT is

23. 23
comprised of a star tracker, three reaction wheels, and the electronics and
software needed to operate the system. ADIOS requires precise pointing and slewing
as the observer continually rotates as it passes the impact cloud and the BCT XACT
is able to provide sufficient attitude control. The reaction wheels in the XACT
provide a slew rate of 10 rad/s with a pointing accuracy of ∓0.003 degrees for 2 axes
and ∓0.007 degrees for the third axis.37 All of the components are housed in an aluminum frame that
helps shield the unit from radiation. This system has flown on the CubeSat missions such as MinXSS and has been
chosen to fly on one of the first interplanetary CubeSat missions, MarCO. This flight pedigree has proven the
system’s reliability and makes it an easy choice for the ADCS of ADIOS.
The BCT XACT was chosen because it contains all the attitude determination and control components that
will work near Deimos. Because Deimos does not have a magnetic field, magnetorquers and magnetometers cannot
be used. Other systems were discarded because they had either attitude control that relies on a magnetic field or
attitude determination that relies on a magnetic field.
(6) GNC
Observer
Delta Differential One-Way Ranging (DDOR) is an interplanetary tracking and navigation technique that
will be used on the Observer for its GNC. NASA’s Deep Space Network (DSN) has been assisting the ESA since
1986 in the guidance and navigation of their missions. The tracking stations for DDOR are located at ESA’s deep
space stations in New Norcia in Western Australia, Cebreros near Madrid, Spain, and Malargüe, Argentina. The
stations provide very accurate measurements.33
The DDOR technique uses two antennas to track a transmitting probe. The two antennas are widely spread
apart and will be tracked simultaneously. By tracking the antennas simultaneously, the time delay between the
arrival of the signals at the two stations (any two of the three ESA stations may be used) can be calculated. In
theory, the time delay should only depend on the positions of the two antennas and the spacecraft.33 However, a
multitude of errors affect the time delay between the two signals. Sources of error may include clock instabilities at
the ground station, solar plasma, and the travel of radio waves through the Earth’s atmosphere, specifically the
troposphere and ionosphere. In order to account for theses errors, DDOR tracks a quasar which is an active galactic
nucleus that emits massive amounts of electromagnetic energy.3 The quasarmust be seen in a direction close enough
to the spacecraft in order for proper calibration to take place. The chosen quasar’s direction is known through
astronomical measurements. The astronomical measurements allow for the quasar’s direction coordinates to be
extremely accurate. With the quasar usually being within ten degrees of the spacecraft, the source signals take a
similar path through Earth’s atmosphere. The time delay of the quasar is subtracted fromthe spacecraft’s time delay
in order to calculate the actual DDOR measurement. The actual measurement is then converted to distance by
multiplying the time delay by the speed of light. The DDOR is valid and verified technique. It has been used on
interplanetary spacecraft such as the Mars and Venus Express. DDOR is expected to be used in future missions as
well.
While GPS is often used for CubeSat missions today, this technique is not available for interplanetary
missions. This is why DDOR was chosen. As the most popular ways to performinterplanetary GNC by the ESA, it
is known to be a very accurate system. Another systemthat was discussed was through the use of sun sensors and
the position of the Sun in respect to the Earth. Given these two, it is possible to triangulate to position of the
CubeSat relative to Earth and the Sun, however, the error of these calculations would be significant.
Impactor
On the impactor, GNC will be performed using the ECAM-M50 Imaging System from Malin Space
Science Systems. This imaging systems consists of a 5 Megapixel Monochrome CMOS camera that can perform
target-relative navigation near Deimos.43 This system was chosen for its small size and performance, shown to be
sufficient as NASA has chosen this platform for its Near-Earth Asteroid Scout (NEA Scout) mission.15 Using the

24. 24
ECAM-M50 in conjunction with the star tracker on the ADCS and image processing software, the GNC systemwill
be able to guide the spacecraft towards Deimos around a day after separation fromthe Observer. This will allow for
sufficient time and distance to make all necessary maneuvers to impact Deimos on time.
Optical navigation has been used on past missions, such as Rosetta, where precise navigation was required
for mission success. In the case of Rosetta, the spacecraft had to rendezvous with a comet and take high resolution
images of the nucleus.18 Like Rosetta, the navigation camera (NAVCAM) will use landmark observations on the
surface that will be obtained through processing NAVCAM images. Landmarks that can be used on Deimos include
the craters Swift and Voltaire. The images would then be made into a 3D map through a process called
stereophotoclinometry where grey levels of multiple images are translated into slopes. The slopes are integrated into
heights and this allows the production of a map. Using these landmarks, the CubeSat can aimfor this area and make
maneuvers to impact this space.
The use of DDOR was considered for GNC of the Impactor, however, the power requirements of an IRIS
communication system were too great to add to the Impactor. Furthermore, the size of the imaging systemprovided
more space for other subsystems. The other systems that were considered for the Observer were also considered for
the impactor, however, since accuracy is crucial for the Impactor these systems were not chosen.
(7) Telecommunications
Observer
One of the toughest architectures to select, based upon the sizing restrictions, was the telecommunication
system. The Iris V2 transponder could satisfy the sizing and power requirements while allowing for direct
communication to Earth over the DSN. The peak power requirements of the transponder is 26 watts.47 Four watts of
which are utilized in transmitting to a 34 m DSN dish on Earth. The rest is lost to heat. Iris will transmit using a Tx
patch 20cm x 20cm, fixed on one face of the Observer. This patch is capable of 1000 bps at opposition and 62 bps
at conjunction.47 It has a 21dBi gain. For receiving, the same panel will have an integrated Rx panel. The Iris V2
system ways 1.2 kg and is 0.5U in volume.47 The Iris V2 is currently TRL 5, but will launch on MARCO 1 and raise
the TRL before the Mars 2020 launch, it is being developed by JPL.2,8
(8) C&DH, Flight Software
Observer C&DH
From the estimation of the maximum data needed to be stored, 5.16 MB, the Cube Computer was selected
for the Observer. It has an operating voltage of 3.3 Volts, a 3.3V and 5V I2C bus voltages, and operates in a
temperature range from -10 degrees Celsius to 70 degrees Celsius.40 It has two 1 MB external SRAM for storage, as
well as MicroSD socket for additional storage up to 2GB.40 The mass ranges from50 to 70 grams depending on how
the motherboard is configured. The power consumption is less than 200 mW.40
Another option is the ISIS On Board Computer (iOBC). It has an operating voltage of 3.3 Volts, a 3.3V
and 5V I2C bus voltages, and operates in a temperature range from -25 degrees Celsius to 65 degrees Celsius.52 For
data storage, it has two slots for either 8 GB high reliability SD cards or any size standard SD cards.52 It also has
optional daughter boards for extra interfaces. The downside to this computer is that it has a mass of 94 grams with
the daughter board and has a power consumption of 400 mW on average.52 Due to the lower maximum mass and
lower power consumption, the Cube Computer was selected.
Impactor C&DH
For the Impactor, the Gomspace NanoMind A3200. It has an operating voltage of 3.3V and I2C interface.60
It operates in -40 to 60 degrees Celsius.60 It has 512 KB build-in flash, two 64 MB NOR flash, 32 kB FRAM, and 32
MB SDRAM.60 Its mass is only 14 grams.60
Flight Software
The flight software will take transmission signals and compute its position. Once the pos ition is calculated,
the ADCS can orient the spacecraft for alignment burns. At a certain position and time, the Observer will activate
the NiChrome wire cutters to release the Impactor. At this point, the Impactor turns on and the Observer rotates to

25. 25
do the retrograde burn. The Impactor uses the camera to determine its position relative to Deimos, and the ADCS
will orient the Impactor for small course correction burns.
(9) Payload Accommodation
To allow the Observer and Impactor to separate from each other, a NiChrome wire cutter mechanism will
be employed. This mechanism works by having a Vectran tie down cable hold back a compression spring system
that will push the Impactor away from the Observer. The vectran cable will be cut by a NiChrome wire when a
constant current is applied to it. This current heats up the NiChrome wire, which in turn thermally cuts the Vectran
cable. Once the Vectran cable is cut, the compression spring is released and pushes against the Impactor.
The average cutting time of 200 and 400 Denier Vectran in a vacuum is 2.6 seconds and 3 seconds
respectively.61 This allows plenty of time for the current in the NiChrome wire to cut no matter the length of
NiChrome chosen. For a 0.5 inch length NiChrome wire, the failure current is 2.15 amps and failure time is 7
seconds.61 For a 3.5 inch length, the failure current is 1.9 amps with a failure time of 19 seconds.61 This gives plenty
of time for the Vectran cable to be cut. For redundancy, secondary circuits with NiChrome wire cutters can be
placed in case the first NiChrome wire cutter circuit fails. Figure 12 below shows where the primary release
mechanisms will be placed, and Figure 13 shows the mechanism model.
Figure 12. Depiction of release mechanism location

26. 26
Figure 13. Release mechanism

27. 27
6.0 System Engineering
A. Technical Resources
(1) Mass Budget
The mass of ADIOS is divided into two modules, the Observer and the Impactor. The Impactor’s mass is
expected to be 4.4 kg with propellant so that when it hits Deimos, it will be slightly under 4 kg. This is the desired
mass, as determined by the plume generation analysis. The Impactor’s dry mass affects the propulsive
maneuverability of the module and was taken into consideration. As for the Observer, the budget does not have to
adhere to the same constraints and the maximum expected value is 9.4 kg, with propellant.
The Observer’s mass is dominated by propulsion and power, followed by ADCS and the communication
system. This is expected, as the Observer is responsible for communication and propulsion for most of the flight.
The rest of the module’s mass is comprised of the structure and other incidentals. As shown in Table XIV, each
component was given a mass contingency based on its TRL. Levels of 8 and 9 were given a 5 percent contingency
while the developing technologies such as the Iris V2 and MPS-130 were appropriately designated higher
contingencies, to match their lower TRL values.
Table XIV. Observer Mass Budget
Subsystem
Component
(Quantity)
Current Best
Estimate (kg)
TRL Contingency (% )
Maximum
Expected Value
(kg)
ADCS BCT XACT 0.91 9 5 0.956
Communications Iris V2 1.2 5 25 1.5
C&DH Cube Computer 0.07 9 5 0.074
EPS
FlexU EPS 0.148 8 5 0.155
60 Wh Battery 0.475 8 5 0.499
2U Deployable
Array (4)
0.8 8 5 0.84
Payload
Argus IR
Spectrometer
0.23 9 5 0.242
Propulsion (Wet) MPS-130 3.5 6 25 4.375
Structure
Aluminum Frame
(2)
0.171 9 5 0.179
Radiation
Shielding
0.25 9 5 0.263
Fasteners (40) 0.2 9 5 0.21
Cables, Wires
(20)
0.1 9 5 0.105
Subtotal (Dry) 6.754 8.098
Subtotal (Wet) 8.054 9.398
The Impactor is mostly comprised of its propulsion, power and ADCS subsystems. This breakdown is
appropriate as the module has no real payload and must impact Deimos to avoid mission failure. The Impactor must
be able to operate on battery power alone while targeting Deimos. As shown in Table XV, the same approach as
before was used to determine the level of contingency allocated to each subsystem.

28. 28
Table XV. Impactor Mass Budget
Subsystem
Component
(Quantity)
Current Best
Estimate (kg)
TRL
Contingency
(% )
Maximum Expected
Value (kg)
ADCS BCT XACT 0.91 9 5 0.956
C&DH NanoMind A3200 0.014 9 5 0.018
EPS
FlexU EPS 0.148 8 5 0.155
40 Wh Battery (3) 0.954 8 5 1.001
GNC MSSS ECAM-M50 0.256 7 20 0.307
Propulsion (Wet) End-Mounted MiPS 0.924 6 25 1.155
Structure
Aluminum Frame 0.401 9 5 0.421
Radiation Shielding 0.15 9 5 0.158
Fasteners (25) 0.125 9 5 0.131
Cables, Wires (15) 0.075 9 5 0.079
Subtotal (Dry) 3.534 3.958
Subtotal (Wet) 3.957 4.381
Table XVI holds the total mass values for the combined Observer and Impactor modules, with and without
propellant. At 13.8 kg, ADIOS meets the mass requirement.
Table XVI. Spacecraft Mass Budget
Total Current Best Estimate (kg) Total Maximum Expected Value (kg)
Dry 10.288 12.056
Wet 12.011 13.779
(2) Power Budget
Since the Observer and Impactor will each have their own power source and power system, separate power
analyses were done for each one.
For the Observer, the power budget shown in Table XVII shows the four operational modes it will operate
in. It is important to note that the EPS power requirement is accounted for in the Solar Cell output as part of the
bundle, thermal control is passive so it does not require power, and GNC power requirement is accounted for in the
communication requirement. Our missions has a very linear timeline so we divided it into different modes. Each
mode will have certain subsystems activated. For the Maneuvering Mode, the structural subsystem(the mechanisms
that release and separate the Impactor from the Observer), the spectrometer, and the downlink communication are
not activated. During the Cruise Mode, the spectrometer, propulsion, and downlink are not activated; however, the
separation mechanism is accounted for in this phase. During the Science Mode, as the data is being collected,
downlink is the only component not activated. Finally, during the Downlink Mode, the spectrometer and uplink
communication are the only components not activated. Overall, the highest average power is during the Maneuver
Mode at 26.06 Watts. For the most of our mission the solar panels will be providing all the power (27.08 W) and the
batteries will not need to be used or charged until the spacecraft is in downlink mode during which time the batteries
will be used. Downlink Mode will require 52.51 W for approximately 10 minutes, so the 60Wh battery will provide
more than enough power. Figure 14 shows the power timeline for our mission.
Table XVII. Observer Power Budget

29. 29
Subsystem
Component Average Maneuver Cruise Science Downlink
CBE
(W)
Cont.
%
MEV
(W)
Duty Cycle
Duty
Cycle
Duty Cycle Duty Cycle
Spectrometer 1.24 15.00 1.43 0 0 1 0
Structure 5.83 20.00 7.00 0 1 0 0
OBP 0.13 5.00 0.14 1 1 1 1
AD&C 0.50 15.00 0.58 1 1 1 1
Propulsion 11.00 5.00 11.55 1 0 1 1
Uplink 12.00 15.00 13.80 1 1 1 0
Downlink 35.00 15.00 40.25 0 0 0 1
Total Power 60.93 26.06 21.51 25.69 52.51
Avg. Power Provided
(At Mars)
27.60
Peak Power Provided
(At Mars)
60.00
Figure 14. Timeline For Observer Power Mode and Power Supply use
For the Impactor, the operation time will be 50 hour and will require 15.68 Watts of power for each
propulsive maneuver, lasting for minutes each time. The selected battery for the Impactor has a 120 Wh capability,
which is more than enough to meet the power requirements and allow the impactor to make as many maneuvers as
necessary. Once again, the EPS power requirement is already accounted for in the battery power output and the
thermal control is passive so it will not require power. Finally, there are no structural power requirements as the
separation mechanism will be powered by the Observer. Table XVIII shows the power budget for the Impactor.

30. 30
Table XVIII. Impactor Power Budget
Average Component Estimated Draw
Subsystem CBE Power (W) Cont. % MEV Power (W)
Structure and Mechanisms 0.00 0.20 0.00
Thermal Control 0.00 0.20 0.00
Power (inc. harness) 0.00 0.10 0.00
On-Board Processing 0.55 0.05 0.58
Attitude Determination and
Control
2.00 0.15 2.30
Propulsion 10.00 0.05 10.50
Guidance and Navigation Control 2.00 0.15 2.30
Total Power 14.55 15.68
Operational Time 1 Hour
Power Provided 40.00
(3) Telecommunications Link Budget
During the hibernation phase, the Observer should only require occasional communication for navigation
and system checks. When separation occurs, it will require more intensive navigation use. During plume analysis
phase, communication systems will be off. Once this ends, the Observer will enter full transmit to relay its data.
There is no time limit to this and it may extend into mission end of life operations.
Table XIX. Downlink Budget
Parameter Value Units
Spacecraft (High Gain, X-
band)
Downlink Frequency 8400 MHz
Antenna Gain -21 dBi
Transmit Power 4 W
Line Loss -4 dB
Transmit bitrate 62 bps
Range
Space Loss -278 dB
Atmosphere Loss -5 dB
Total Path Loss -283 dB
Ground Station
Dish Diameter 34 m
Antenna Efficiency 90 %
Gain 69.1 dB
System Noise Temp 14 K
Link Margin
Available Eb/No 9.2 dB
Modem Loss -1.2 dB
Required Eb/No 6 dB
Link Margin 2 dB

31. 31
Analysis of the link budget utilizing the Iris V2 transponder has proven to be difficult. However, based
upon current estimates of hardware capabilities, the link will close.
(4) Thermal Energy Balance Assessment
Table XX. Observer Operational Temperatures
Observer Component
Lowest
Temperature (°C)
Maximum
Temperature (°C)
ADCS BCT XACT -40 80
Communications IRIS -20 50
C&DH Cube Computer -10 70
EPS
Clyde Space FlexU EPS -40 60
Clyde Space 60Wh Battery -40 60
Clyde Space 2U Deployable Array -40 85
Payload Argus 1000 IR Spectrometer -20 40
Propulsion Rocketdyne MPS-130 5 50
Structure
Aluminum Frame & Radiation Shielding N/A N/A
Fasteners N/A N/A
Misc. Cables, Wires N/A N/A
Table XXI. Impactor Operational Temperatures
Impactor Component Lowest Temperature (°C)
Maximum Temperature
(°C)
ADCS BCT XACT -40 80
C&DH Nanomind A3200 -25 65
EPS
Clyde Space FlexU EPS -40 60
Clyde Space 40Wh Battery -40 60
GNC MSSS ECAM-M50 -30 40
Payload Ballast N/A N/A
Propulsion VACCO End-Mounted MiPS 0 60
Misc. Cables, Wires N/A N/A
Structure
Fasteners N/A N/A
Aluminum Frame & Radiation N/A N/A

32. 32
Shielding
𝐵 = √((𝐵 ∙ 𝐵 𝐵 )/(𝐵 ∙ 𝐵 ∙ 𝐵))
4
(11)
Assessing the crafts thermal needs, the equation above was used to determine the internal temperature of
the cubesat. The craft stays thermally constant for the most part since it is only out of the sun’s line of sight for very
short amounts of time. The maximum and minimum allowable temperatures are determined from the tables above.
The components with the lowest absolute value for temperature are what the tolerances are tailored around. The
Observer experiences maximum and minimum temperatures of 28 and 18 degrees Celsius, respectively, while the
Impactor’s maximum and minimum temperatures are 30 and 21 degrees Celsius, respectively. These fall into the
acceptable range passively meaning no thermal control is required for either the Observer or Impactor.
Radiation is a serious topic for interplanetary missions. According to measurements taken by Curiosity on
its transit to Mars, it experienced 300 mSv over 180 days. That is approximately 1.67 mSv/day and for a 210 day
journey that results in a total of 350 mSv of radiation that our CubeSat will experience. All of the components on the
CubeSat can withstand 3 years of radiation dosage in LEO. One month in LEO amounts to 80 mSv on average or
0.44 mSv/day. This means that the total dosage the CubeSat can withs tand is about 480 mSv which is over 35% the
expected amount it will receive throughout the mission. To detect and reduce the effects of single-event upsets, a
watchdog processor will be utilized is a hardware timer that automatically generates a system res et if the main
program neglects to periodically service it.
(5) Data Return Strategy
Iris V2 will use X-band to transmit data back. With the requirement of relaying 5.16 MB is more than
feasible, as ADIOS will have the remainder of its mission to do this. At peak transmission rates this will only take
10 minutes. At minimum transmission rates this will take 184.9 hours, or 7.7 days. During this time the transmitter
will maintain a Eb/No necessary to close. If for some reason the trajectory doesn’t allow for closed link, then the
observer will come back into range after a few months.
(6) ΔV and Propellant Budget
The total amount of ΔV required varies greatly by launch day, with later launch/separation dates reducing
the total ΔV required. The largest contributor to the overall ΔV is the final burn, ΔVo. While it is the least varied
among the burns by date, only about 25 m/s, it accounts for about half of the ΔV early on. Both ΔVc and ΔVi vary
by nearly 50 m/s overall which makes the later launch dates even more desireable as it creates 2 large ΔV
reductions.

33. 33
Figure 15. ΔV Requirements by Launch Date
The average ΔV values and the differences from the maximum values can be seen in Table XXIII. The
average total ΔV required is only 136 m/s but can be as high as 219 m/s or as low as 96 m/s. That’s up to a 61%
maximum margin and a -30% minimum margin. While this is a large range, it can be determined pre-launch and so
the amount of fuel loaded can be adjusted to meet the overall mission requirements. The variance columns indicate
the difference of the ΔV value from the average. In the case of ΔVi, while the average burn is 41.46 m/s, the
maximum burn will be 14.781 m/s greater and the minimum burn will be 26.4 m/s less than the average.
Table XXII. Variation of ΔV Requirements
ΔVi Variance ΔVc Variance ΔVo Variance ΔVT Variance
Average Value 41.46 0 19.00 0 75.17 0 135.64 0
Maximum
Value
56.24 14.781 62.63 43.63 104.67 29.50 218.96 83.32
Minimum
Value
15.06 -26.40 8.19 -10.81 67.24 -7.94 96.12 -39.52
The impact velocity does not vary greatly with the amount of ΔV required for the mission. This means
regardless of launch date we will still be able to acquire a sufficient impact velocity for the plume to develop. After
looking through all the data, the most ideal launch dates would be after the July 20th maximum for the ΔVi burn.
While ΔVo and ΔVc do increase again after the 27th, ΔVi’s reduction compensates enough to warrant launching
then.

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Figure 16. Impact Velocity as a function of Total ΔV Required
(7) Margins Assessment
Individual margins are discussed in their respective sections. The overview for the margins can be seen in
Table XXIV below.
Table XXIII. Margin Assessment
Mission
Requirements
Spacecraft
Characteristics
Margin (% )
Mass (kg) 14 13.777 1.6
Power (Observer) (W) 52.5 60 14.3
Power (Impactor) (W) 15.7 40 155
∆V (m/s) 218 233.3 7.01
Propellant Mass (Observer) (kg) 0.95 1.3 36.8
Thermal (Minimum) (℃) 5 20 5.39
Thermal (Maximum) (℃) 40 30 3.19
Below is the margin for plume size for different material properties, from loose sand to solid rock, and
impact speeds. The red dot indicates the characteristics the mission was designed to. The purpose of this graph is to
clearly show that for almost all changes in variables the plume is far larger than designed for and always larger than
the minimum of 2 km in height.

35. 35

36. 36
7.0 Risk Identification & Mitigation
Technical Risk Areas and Mitigations
Figure 17. Mission Risk Matrix
This cubesat mission is of moderate risk.
1. Damage to key systems from radiation
2. Trajectory Mishap
3. Impactor Separation Failure
4. Plume Size Failure
5. Power Failure
1. Damage from radiation could potentially cause any of the main or secondary instruments to fail or malfunction.
Average radiation dose in interplanetary space is approximately 300 mSv for a 180 day journey which is 3 times
what is experienced in LEO. To mitigate this risk all components of the cubesat were chosen for hardness against
radiation.
2. A trajectory perturbation could cause the mission to arrive at Mars at an inopportune time or even miss Deimos all
together. In the event of a successful burn by Mars 2020 ADIOS but a late launch date, ADIOS has enough ΔVto
arrive at Mars at an optimal time for all times during the launch window. In the event of a trajectory perturbation
during the burn by Mars 2020, ADIOS has 33% extra ΔV on top of the worst case scenario launch date. Anything
more than this would probably result due to the failure of the Mars 2020 burn and would not be able to be mitigated.
3. If the Impactor fails separation the mission would have to be reduced to a surface spectronomy mission. To
mitigate this, a NiChrome wire cutter was chosen to be our separation mechanism. It it a reliable and simple system
that reduces the changes of Impactor separation failure.

37. 37
4. If the plume is not big enough due to unforeseen variables, the spectrometer might not be able to point at it
resulting in a mission failure. This is managed by the Impactor’s technical specifications being calculated at a worst
case scenario level. Any change in variable will most likely result in a larger than expected plume.
5. A power failure would mostly likely be mission ending. At the time of spectronomy operations, almost no system
can be powered down to save power. To mitigate this, a 4 watt hour battery in the observer is included to both deal
with peak loads and possible solar power failure.
Project-level Risk Assessment and Mitigations
1. Propulsion isn’t ready in time
2. Production schedule surpasses time allotted until launch
3. Components don’t meet testing standards
4. Costs are higher than expected
5. Loss of satellite upon launch
Figure 18. Project-level risk assessment matrix
There are five major project-level risks to be considered in risk analysis, the likelihood and consequences of which
are summarized in Figure 18.
1. An overdue in propulsion systemdelivery is mission-critical, particularly for the Rocketdyne MPS-130
propulsion module. The high delta-V capability is essentialfor the success ofADIOS mission; however, the MPS-
130 has been in test phase for a long time without anticipated date of availability. The VACCO cold gas propulsion
module, on the other hand,has samples already available and is generally higher in TRL than the MPS-130. The
cold gas propulsion is also less mission critical and can be replaced by other similar products should the delivery fail
to achieve.
2. An overdue in spacecraft production also leads to direct failure of mission, since the launch vehicle and launch
date is not under control of the ADIOS mission team. The likelihood of an project overdue is to the best reduced by
setting the end date of Phase D (systemassembly, integration and testing) at least sixmonths prior to launch, giving
a generous margin for any delay.

38. 38
3. Unsatisfactory quality of spacecraft components threatens the reliability of spacecraft and thus the possibility of
mission success. However, many of the components are off-the-shelf and have been tested by other successful
missions. The component in development, such as the MPS-130, will be manufactured by top-tier companies, and
thus the chances of component failing standards are low. Other structural parts, including frame, wire and
thermal/radiation protection are held to relatively low standards, given the somewhat short (less than one year)
lifespan of the spacecraft. In addition, redundant testing will be performed to guarantee readiness of all components;
extra testing and possible fixing of parts are accounted for in schedule design.
4. A cost overdue is likely, given the unpredictability of project progress. For instance, increase of cost will occur if
components are replaced, duplicates are used for testing, or if salary level increases over the year at a rate higher
than anticipated. However, reasonable cost overdue are covered in contingency and will not be impeding the
mission. Should cost rise above the maximum anticipated level, it is most likely still within the $5.3M limit and thus
additional funding can be acquired.
5. A loss of satellite upon launch will lead to direct failure of mission, but the possibility is low given the very high
reliability of launch vehicle (Atlas V), with only one partial failure per 66 flights. The sub-model Atlas V-541 used
for Mars 2020 mission has three successful launches to date with no failure. Should a launch vehicle accident occur
and the spacecraft is lost, cost would be reimbursed partially by insurance.

39. 39
8.0 Management, Schedule & Cost
Figure 19. Work Breakdown Structure
A. Management Plan
The mission management plan will allocate positions and tasks to each team member. The WBS shown
above will be used to organize the project roles and tasks. One graduate student will be assigned to each manager
position, giving a total of five graduate students. Two undergraduate students will be assigned to each supervisor
position, giving a total of eighteen undergraduate students. There will be one principal investigator for this mission,
giving the team a total of twenty four members.
The mission program manager will be the primary contact between the team, the principal investigator, and
NASA offices. The mission program manager will be responsible for all non-engineering related tasks. This will
include, but will not be limited to travel arrangements, cost budgeting, setting up meeting times and deadlines,
making sure each subteam and the entire team is on track, and settling disputes within the team. The mission
program manager will have final decision authority throughout the entire mission.
The mission operations manager will be responsible for making sure all operations and mission designs are
viable and on track. The flight operations manager will be responsible for project management, flight dynamics, and
payload operations. The project management supervisors will work closely with the mission program manager in
order to develop an efficient schedule and method of conducting meetings for each subteam and the entire teamas
well. Facilities and administration for testing and integration will also be handled by the project management
supervisors and the mission program manager. The flight dynamics supervisors are responsible for trajectory,
guidance and navigation, attitude determination and control, and propulsion. Selection of the propulsion system as
well as determining the mission schedule will be key tasks. The payload operations supervisors will be assisting the
flight dynamics supervisors. Once the payload has been selected, there is not much to be done until the testing and
integration phases begin. The launch operations manager will oversee the launch preparation and post-launch
operations. Launch preparations will involve degassing and the packaging and transportation of all necessary
equipment/components for the mission. Post-launch operations will consist of analyzing all data that will be
collected throughout the duration of the mission.

40. 40
The mission engineering manager will be responsible for physical and electrical systems as well as testing
and integration. The physical system supervisors will be responsible for developing the separation mechanism
between the observer and the impactor. Thermal and radiation analysis in order to ensure survivability will be
performed by the physical system supervisors. The electrical systems supervisors will be responsible for the
electrical power systems of the observer and impactor. A power budget will also be created in order to ensure that
the mission stays within the power requirements. Communications, command and data handling, and the telecomm
budget will also be completed by the electrical systems supervisors. The integration supervisors will be responsible
for the mass and size budgets, as well as developing CAD models for the mission. The testing supervisor will be
responsible for thermal, radiation, and vibration testing. The integration and testing supervisors will work closely
with both systems supervisors, the integration supervisors and the flight dynamics supervisor. This is necessary in
order to ensure that the entire team is on the same page. Miscommunication amongst team members can result in
mission-fatal mistakes.
It is to be noted that despite the teambeing separated into their own specific roles, collaboration within the
team is necessary for mission success. For example, the payload operations supervisor must make sure that the
spectrometer selected meets mass, size, and power requirements. If not, the electrical systems and integration
supervisors must work with the payload operations supervisors in order to come up with an alternate solution.
B. Program Schedule
Figure 20. Program Schedule
Pre-Phase A
Pre-Phase A will consist of concept studies relating to the mission. This may include refining the mission concept,
preliminary mission architecture design, and assessing technology and engineering development. Using these, staff
and infrastructure requirements will be determined for the project. With the work done in Pre-Phase A, a
Formulation Agreement Document (FAD) will be developed as a tool for communicating and negotiatin g the
project’s funding and schedule requirements for mission development. Before entering Phase A of the project, a
Mission Concept Review (MCR) will be conducted to ensure life-cycle review objectives and expected maturity
states have been met.52
Pre-Phase A will last approximately 2 months from January 2017 to the end of February 2017.
Phase A
Phase A is the first phase of project formulation. Within this step, concept and technology development activities

41. 41
will be performed to develop a mission architecture that meets programrequirements and constraints for the project.
This will ensure the mission definition and plans are mature enough to begin Phase B.
The activities in this phase include:
• Developing/defining project requirements to the system level
• Developing system architecture
• Developing cost and schedule estimates for the project
• Identifying and mitigating development risks
• Updating staffing/infrastructure requirements
In this phase, two reviews will be conducted. In the middle of this phase, a System Requirements Review (SRR)
will be performed to evaluate whether the performance requirements of the system satisfy the program’s
requirements of the project. At the end of Phase A, a Mission Definition Review (MDR) will be conducted to
determine if the maturity of the mission definition is sufficient to begin Phase B.52
Phase A will last 5 months from March 2017 to the end of July 2017.
Phase B
Phase B is the final phase of project formulation and deals with the preliminary design and technology completion
activities. In this phase, technology development will be completed, heritage hardware and software assessments
will be performed, engineering prototyping will be conducted, and risk-mitigation activities will be identified.
Furthermore, mission objectives, requirements, cost estimates, and schedule estimates will be updated.
At the end of this phase, a Preliminary Design Review (PDR) will be conducted. The PDR will consist of
evaluating the planning, cost, and schedule developed during Formulation and assessing the preliminary design’s
compliance with the mission requirements.52
Phase B will last approximately 8 months from August 2017 to the end of March 2018.
Phases C-D
Phases C and D are the steps in which project implementation is conducted. In Phase C, final systemdesigns will be
completed and documented, and fabrication of test and flight architecture will begin. Phase D consists of system
assembly, integration, and testing various systempieces, while performing verification and validation on products as
they are integrated. During this phase, hardware and software documentation will be finalized, issues from testing
will be resolved, and the systemwill be prepared for launch and shipment.
In Phase C, two reviews will be conducted, the Critical Design Review (CDR) and the System Integration Review
(SIR). The objective of the CDR is to evaluate the integrity of the project design and the ability of the system to
meet mission requirements with appropriate margin and risk.52 The CDR is performed in the middle of Phase C.
The SIR will evaluate the readiness of the project to begin assembly, integration, and testing. The SIR is performed
at the end of Phase C.52
In Phase D, the Operational Readiness Review (ORR) and Mission Readiness Review (MRR) will be conducted.
The ORR is used to evaluate the readiness of the project to assemble, integrate, and test flight systems.52 The ORR
is performed in the middle of Phase D. The MRR is used to assess the pre-flight operational readiness of the
project.52 This will insure there are no problems with the flight systems and all personnel and infrastructure are
ready to support launch. The MRR is performed at the end of Phase D.

42. 42
Phase C will last around 15 months (from April 2018 to the beginning of July 2019) and Phase D will last around 11
months (from July 2019 to the beginning of July 2020).
Phase E-F
Phase E is the launch, operations, and sustainment of the mission. The parameters for this phase are determined by
the flight schedule. In this phase, Decommissioning Review (DR) is performed. The DR (towards the end o f the
mission) will evaluate the readiness of the project to conduct closeout activities.52
Phase F deals with the decommissioning of the mission. This phase begins with a Disposal Readiness Review
(DRR), in which the readiness of the project for disposal will be evaluated.52 In this stage, final archival of data and
spacecraft closeout will be performed.
Phase E will last under 9 months (from the beginning of July 2020 to the end of March 2021) and Phase F will last 2
months (from April 2021 to the beginning of June 2021).
Figure 21. Critical Path
Critical Path
The critical path of the project shows that the minimum duration of the project will be 50 months. Each of the
activity times are broken down in the Phase summaries above. This path does not include schedule margin, which is
discussed below.
Schedule Margin
The planned activities of this mission will last 52 months, providing 8 months of schedule margin. This will provide

43. 43
sufficient margin for activities that may take more time than expected such as assembly, integration, and testing.
The schedule margin rate is 0.133.
C. Cost Estimate
Figure 22. Mission Cost Budget
The cost estimation for the mission will be performed using the Georgia Tech Office of Sponsored
Programs (GTOSP) project proposal template.43 Costs from the year 2017 to the year 2021 will be estimated in order
to ensure that the mission is monetarily feasible. The main methodology will be to directly account for each
individual/component necessary for mission success. Figure 22 details the overall mission cost budget.
Personnel
The R&D team of ADIOS consists of one principal investigator (PI), five graduate students, and eighteen
undergraduate students, giving the team a total of twenty four members. Salaries and benefits will be calculated by
directly accounting for each individual.The PI will require a base salary of $113,381.00 (FY17) with a 3% raise each
year. The graduate students will require a base salary of $26,000 each (FY17) with a 1.5% increase each year. Only
the PI will receive fringe benefits (28.8% of annual salary) due to the fact that all other team members are
considered to be part-time. The total cost of salaries and wages for the entire duration of the project came out to be

44. 44
$1,445,112.90. Roughly 38% of the total cost will be towards personnel. Annual salaries as well as salary increase
and fringe benefit rates were obtained from GTOSP.56
Equipment
Each necessary component was directly accounted for by directly contacting the manufacturer (excluding
the VACCO End-Mounted MiPS and Rocketdyne MPS-130). The cost of the two propulsion systems, the VACCO
MiPS and the MPS-130, have not been disclosed by manufacturers, and will be estimated based off similar products.
The two propulsion systems will be given a high cost contingency (40%) in order to emphasize the uncertainty of
the prices.. Extra solar panels, batteries, and computer chips will be purchased in order to account for the possibility
of equipment breaking during testing/integration. $30,000 for materials and supplies will be allocated for each year
for any supplies required throughout the duration of the mission. All necessary equipment will be p urchased over a
three year span (2017-2019). This is due to the fact that purchasing all the equipment in one year is monetarily
unfeasible. The total cost of equipment came out to be $801,416.00. Cost contingency for each component was
individually assigned based on availability and TRL. Taking contingency into account raised the cost of the mission
by $181,298.81.
Travel
Travel was estimated by selecting locations for a design review trip and an aerospace conference. Living
and transportation arrangements were estimated for six people (PI and 5 managers). Locations were selected to be
Pasadena, California and Big Sky, Montana (design review trip at JPL and 2017 IEEE Aerospace Conference
respectively).46 Both trips will last approximately one week. The total cost of travel over the entire duration of the
project will be $129,576.05.
Other Direct Costs
Other direct costs include documentation/publication, graduate student tuition and healthcare, and Deep
Space Network cost. The cost of documentation/publication is a one-time payment of $5,000.28 For graduate
students, health insurance and tuition were estimated using rates obtained fromGTOSP.43 Healthcare insurance rates
will be set at 4.7% (FY17) of the annual salary. The tuition rate will be set at 5% (FY17) of the annual salary per
month. Healthcare and tuition for graduate students totaled to be $433,357.17. The Deep Space Network cost will b e
calculated using an equation provided by NASA.32 The equation calculates aperture fees for DSN. The cost for five
years (inflation is accounted for) came out to be $12,949.32.
AF = RB [AW (0.9 + Fc/10)] (12)
Indirect Costs
Indirect costs will be calculated using rates obtained from GTOSP.43 57.8% of the direct costs excluding
graduate student tuition and equipment will be the indirect costs associated with the mission. The indirect costs over
a five year span came out to be $1,018,344.09.
Total Cost
The total cost of the mission, over a five year period, came out to be $3,995,755.54. Taking contingency
into account, the cost increased to $4,177,054.35. Both total costs fall well below the budget limit of $5.3 million.
The cost in terms of WBS elements is detailed in figure 23, below. The majority of the costs comes from project
management. This contains all direct costs as well as all overhead costs. Besides project management cost, the rest
of the mission’s cost comes from the equipment required for the mission.

45. 45
Figure 23. Cost Breakdown
D. Descope Options
There are several options available for descoping. Reducing the amount of testing performed on each
component will reduce the amount of time necessary for testing. This reduces the overall required build. Reducing
the amount of employees or the amount of trips taken serve as cost descope options. By removing two graduate
students, most likely the flight and launch operations managers, costs will be saved in salaries, graduate student
tuition and healthcare as well as overhead costs. Over the five year span, removing two graduate students fromthe
team will save a total of $756,716.35. Other cost descope options include purchasing only one quantity of each item
and hoping that none of the equipment breaks during testing/integration. For example, currently two spectrometers
are set to be purchased, but by only purchasing one spectrometer, a total of $49,500 would be saved. Each
component must be reviewed on an individual basis in order to determine whether purchasing a single quantity will
suffice.
With less testing performed on components, the risk of component failure increases. Reducing the amount
of employees will increase the workload for all other team members. With less employees, there is less collaboration
which reduces the overall efficiency of all phases. Reducing the amount of trips taken will result in less feedback
and critique from experienced professionals. There is also less mission exposure as a result of the reduced amount of
trips.

46. 46
Appendix
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Nomenclature
Aw Deep Space Network Aperture Weighting
ADCS Attitude Determination and Control System
AF Deep Space Network Aperture Fee
CAD Computer-Aided Design
CD&H Command and Data Handling
CDR Critical Design Review
DDOR Delta Differential One-Way Ranging
DR Decommissioning Review
DRR Disposal Readiness Review
DSN Deep Space Network
Eb/No Energy per Bit to Noise Power Spectral Density Ratio
EOL End of Life

49. 49
EPS Electric Power System
ESA European Space Agency
g (FROM EQN 7, PAGE 15)
g0 Gravitational Constant of Earth, 9.81 m/s2
Fc Number of Station Contacts per Calendar Week
FAD Formulation Agreement Document
FOV Field of View
GMAT General Mission Analysis Tool
GNC Guidance, Navigation and Control
GPIM Green Propellant Infusion Mission
GTSOP Georgia Tech Office of Sponsored Projects
Isp Specific Impulse, seconds
JPL Jet Propulsion Laboratory
LEO Low-Earth Orbit
MCR Mission Concept Review
MDR Mission Definition Review
MRR Mission Readiness Review
m0 Initial Total Mass,kg
mf Final Total Mass,kg
NAVCAM Navigation Camera
ORR Operational Readiness Review
PDR Preliminary Design Review
PI Principal Investigator
rp Periapsis Radius
R (FROM EQN 7, PAGE 15)
Rb DSN Contact Dependent Hourly Rate, Adjusted Annually ($1,190.97 per hour, FY16)
R&D Research and Development
SIR System Integration Review
SOI Sphere of Influence
SRR Systems Requirements Review
TCM Trajectory Correction Maneuver
TRL Technology Readiness Level
WBS Work Breakdown Structure
∆V Velocity change, m/s
∆Vi Velocity change for initial burn, m/s
ΔVc Velocity change for impactor burn, m/s
ΔVo Velocity change for observer burn, m/s
ΔVT Total velocity change, m/s