A Deimos Impact and Observation Spacecraft This paper details the initial mission concept proposed for project A.D.I.O.S.

1. Mission Concept
DIOS MIO - Deimos Impact & Observation Spacecraft: Mars Intercept Orbit
Team 3
Jeff Anderson, Thomas Blachman, Andrew Fallon, John Franklin, Samuel Gaultney, David Habashy,
Brian Hardie, Brandon Hing, Zujia Huang, Sung Kim, Jonathan Saenger
Overview/Objectives
Mission Objectives
The main objective of our mission is to analyze the subsurface composition of mars’s moon
Deimos utilizing a spectrometer during a flyby analysis of a debris cloud created by an artificial impactor.
As a secondary objective and a descope the objective of taking close up pictures of Deimos has also been
included. The motivation for this mission come from the decadal survey;
“...There has been little mention of missions to the martian moons, Phobos and Deimos. It is
clear that these two moonscould play important roles in the future exploration of Mars, especially if they
turn out to be related to volatile-rich asteroids, a possibility which has not been excluded by existing data
and observations.If so,they may be the surviving representativesof a family of bodies that originated in
the outer asteroid belt or further, and reached the inner solar systemto deliver volatiles and organics to
the accreting terrestrial Planets….measurements of bulk properties and internal structure, high-
resolution imaging of surface morphology and spectral properties, and measurements of elemental and
mineral composition.”
Mission Overview

2. Utilizing a 6U cubesat and the Mars 2020 trajectory, we plan to release an impactor into Deimos
to release a large dust plume. This plume will be analysed during a flyby with the spectrometer on board
our cubesat. In addition, our cubesat will take images of Deimos to improve the understanding of it’s
geography. This will also serve as our descope Both of these objectives fit well with the desired science
objectives of the Decadal survey.
The driving requirements for our mission revolve around creating a plume large enough for
analysis and likewise being able to analyse a plume to determine the composition of Deimos. This affects
the size of our cubesat because the impactor must be sufficiently large as to have enough kinetic energy at
the velocities the probe is traveling, while keeping the total mass within the requirements. The
spectrometer on the cubesat also must be sufficiently powerful to measure the composition of the plume,
while also being small and light enough to still meet the mission requirement.
Trajectory
The initial trajectory of the Mars 2020 craft has not been fully disclosed yet so a few assumptions
were made. Based on the official Mars 2020 Mission Concept powerpoint (12), the transfer orbit was
approximated and slightly altered to provide a good approach to rendezvous with Deimos. There are two
options for encountering Deimos: prograde or retrograde. Figure 1 shows the approximate approach for a
prograde encounter and Figure 2 shows a retrograde approach.
Fig. 1. Prograde Fig. 2. Retrograde
In the event of being off-phase with Deimos, a ΔV will be performed at Mars’ sphere of influence
(5.9x105
km away) to either slow down or speed up to converge with the moon. The maximum ΔV
required to get back in phase from a maximum phase shift is 257 m/s. That is achievable by our cubesat.
To calculate the speed at which our impactor will strike Deimos, the velocity of the spacecraft, Deimos,
and any ΔV will need to be accounted for. In both cases, the velocity of the spacecraft is 2.4 km/s thanks
to the gravitational pull from Mars, resulting in 3.75 km/s for retrograde and 1.05 km/s for prograde.
These calculations were determined from GMAT. The difference in these two impact velocities greatly
influences how effective the impactor will be. Kinetic energy is proportional to velocity squared, so
nearly quadrupling the velocity from a prograde to a retrograde encounter results in over 12 times the
impact force. This would be more effective by providing a bigger dust cloud to analyze, but it would also
require much more precision regarding trajectory and pointing to collect all the data while travelling so
quickly.
Instruments:
Morphological Matrix

3. -
Architecture
Decision
Option 1 Option 2 Option 3 Option 4
Communication
EWC 27
HDR-TM X
Band
Transmitter
MarCO
Spectrometer
Compact Ion
and Neutral
Mass
COSIMA
Amptek X-
123SDD
Propulsion System Chemical Electric Cold Gas Hybrid
Chemical Hydrazine AF-M315E
Attitude Control
Reaction
Wheels
Cold Gas Electrospray
Attitude
Determination
Star Trackers IMU Sun Sensors
Command & Data
Handling
iOBC
Cube
Computer
-
Impactor
The feasibility of the impactor was calculated using data of the Deep impact mission which
impacted comet tempel-1. Deep impact, impacted with a velocity of 10 m/s and a mass of 300 kg and
ejected 1.2*106
kg of material. Our impactor will hit with a 1/1000 of the kinetic energy and eject 1/1000
of the mass. This is still however 2550 kg of material. We believe this will be sufficient to analyse as it
will make a 2.5 km3
cloud at a density of 1 kg/m3.(4)
The impactor is heavily driven off of the driving requirement of having a plume large enough to
be easily pointed at yet still dense enough to be analysed by the spectrometer. The mass of 3 kg was
chosen in conjunction with trajectory to generate enough impact force as the force scales linearly with the
mass. During the last portion of the flight the impactor will fly independently of the cubesat therefore will
require its own short term propulsion, power, and ADCS, these are listed below.
Spectrometer
The spectrometer selected for our mission to Deimos was the X-123SDD complete x-ray
spectrometer with silicon drift detector (SDD). The X-123SDD spectrometer was selected due to its small
size (7 x 10 x 2.5 cm), low power requirement of 2.5 W, light weight of 180 grams, and a transmission

4. rate of 9.6 kb/s. The Compact Ion and Neutral Mass Spectrometer was not selected because of its larger
weight (560 grams), about three times as heavy although it did have a smaller power requirement. For this
decision we decided to sacrifice .9 Watts of power for a lighter spectrometer. The COSIMA was not
selected because of it required 20.4 W which was more power then we could supply for the spectrometer
alone. The COSIMA was also much larger in size with the following dimensions 39.4 x 97.3 x 37.8 cm.
The X-123SDD complete x-ray spectrometer’s price, set by Amptek incorporated, is approximately
41,600 US dollars. It’s is currently being applied on the Miniature X-ray Solar Spectrometer mission
launched on December 6, 2015.
Propulsion
The main propulsion system for this mission will require a large ∆V
for a CubeSat mission. Initial calculations project this value to be
approximately 250 m/s. The nature of the mission confines the orbital
maneuvers to high thrust in relatively short time periods. A high degree of
mobility will be needed to ensure the CubeSat can successfully position
itself within range to collect data. This constraint suggests the
elimination of electric and solar sail as feasible propulsive methods, as
they require long burn times to achieve high ∆Vs. This leaves two options:
chemical and cold gas. Inert gas propulsion has the advantage of simplicity
but with an average specific impulse of 70 seconds, will not produce a
sufficient ∆V with the given 6U volume constraint. This leaves chemical
propulsion as the best option for the mission.
Bipropellant systems are too mechanically complex to scale down to a CubeSat level, therefore a
monopropellant system will be used. The propellant known as AF-M315E was selected over the
hydrazine for several reasons. Most importantly, AF-M315E has a significantly lower toxicity level when
compared to hydrazine. 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. Additionally, AF-M315 is 45 percent denser than hydrazine which allows for smaller propellant
tanks. It also has a lower freezing point than hydrazine which reduces the complexity of thermal
management subsystem. This propellant has an expected specific impulse of 240 seconds.
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 have
a wet mass of 3.5 kg. The amount of propellant it will be able to hold is 1.3
kg resulting in a 2.2 kg dry mass. Assuming all other systems on the
spacecraft will total 5.5 kg, the maximum expected ∆V is 367 m/s for this
configuration. This is an ample amount for the mission and will allow for 40
percent margin, in case emergency maneuvers are required.
This system has a technology readiness level of 6 and will be flight tested in early 2017 when
NASA’s Green Propellant Infusion Mission (GPIM) launches. This mission is intended to demonstrate
the viability of using AF-M315E as an effective propellant for orbital maneuvers.
System Operations
Attitude Determination and Controls
The primary method of attitude control and determination on the CubeSat will be through the use
of thrusters and the Blue Canyon Technologies (BCT) XACT Attitude Control System. This system was
determined to be the optimal attitude control system through a trade study consisting of four different
systems (as shown in Table I) and comparisons between different attitude control and determination

5. systems in the morphological matrix. While not the cheapest system, it provides the pointing accuracy
and components that are the most optimal for our mission design. The BCT XACT system contains Star-
Trackers,three Reaction Wheels, and the electronics and software to run the system. Since Deimos has no
magnetic field, the Magnetometer and Torque Rods that are part of the MAI-400 system are useless and
the ADS will use star trackers as they will be more accurate than sun sensors. Also, since we are not near
Earth, the IR Earth Horizon Sensors on the MAI-400 will not be useful. One of the biggest perks of the
BCT XACT system is that it is currently being flown on the MinXSS cubesat mission and has been
chosen to fly on one of the first interplanetary cubesat missions, MarCO. The flight pedigree of this
control system makes this the easy choice for our primary ADC system. The best alternative is the BCT
XACT-50 due to its similar specifications and components.
Table I. Attitude Control System for the CubeSat
BCT XACT BCT XACT Lite BCT XACT-50 MAI-400
Pointing Accuracy
(deg)
±0.003, ±0.007 1 ±0.003, ±0.007 N/A
Lifetime (Years) 3 3 3 N/A
Mass (kg) 0.91 0.7 1.225 0.694
Volume (cm^3) 10x10x5 10x10x5 10x10x7.54 10x10x5.16
Electronics
Voltage (V)
5 5 5 5
Reaction Wheel
Voltage (V)
12 12 12 N/A
Slew Rate (deg/s) 10 10 10 N/A
Cost (FY17) $145,123 N/A N/A $72,559
Flights MarCO, MinXSS N/A N/A N/A
Contains Star-Trackers Sun Sensors Star-Trackers 3 Reaction Wheels
Reaction Wheels Magnetometers Reaction Wheels Magnetometer
Electronics IMU Torque Rods 2 IR Earth Horizon
Sensor
Software Reaction Wheels Electronics 3 Torque Rods

6. Electronics Software Computer
Software
The methods of attitude control and determination on the impactor utilize some of the same
components as the main cubesat, however, will be reduced to a smaller package. While the BCT XACT
system is 0.5 U, to cut back on size we will focus on the use of a star tracker, an inertial measurement unit
(IMU), and three reaction wheels to allow for 3-axis control. This will allow for enough attitude control
even with the omission of a propulsion system. As the pointing of requirements of the impactor are less
precise than that of the cubesat, all of the components of the BCT XACT system are not needed and this
can also provide a cost savings. Much like the CubeSat, the use of magnetometers and torque rods are
unnecessary, so only ADCS components that do not rely on magnetic fields can be used. The reaction
wheels that were chosen (Blue Canyon Technologies RWP050) will have a maximum torque of about
0.007 Nm with a power at full momentum of 1 W. The design life is 5 years. The star tracker (Blue
Canyon Technologies Thin Slice NST) will have an attitude solution of 5 Hz, Sky Coverage of over 99%,
and peak power of less than 1.5 W. The IMU chosen is the SMG Ellipse-A with 0.2 degrees of roll and
pitch accuracy over 360 degrees.
Power
The CubeSat will rely solely on Solar panels to generate power. It will use two Clyde Space
Ukube-1 Single 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 of Orbital Power. The power system will be managed by a Clyde Space XU Cubesat
EPS, which is up to 98% efficient regulating 3.3V and 5V Busses. Also, the cubesat will use a 12V
18650B Lithium Ion Batteries with custom battery protection circuitry similar to the one used in MarCo.
For the impactor a 3000mah battery will be used over solar. The short time that it needs to be
used couple with the small size of the impactor makes solar infeasible.
Communication
There are two viable and very different approaches available in terms of communication, the first
is mimicking the transmission system for the MarCO cubesat, launching in 2018. It utilizes an X-band
transmission at 8 kb/sec, it is large and requires far more power than option two, around 20 W, but it will
allow bypass of MRO and direct downlinking to DSN on Earth. Option two is the EWC 27 HDR-TM X-
Band Transmitter, developed by SyrLinks. It has a customizable downlink speed, 5 Gb/s the option for
the 3 m dish on MRO. It requires 10 W of power, and is far smaller and lighter than option one. With the
downlink requirement of 10 kb/s, option two is a superior choice. It is also TL9 mission capable.
Command and Data Handling
For the control of the data and execution of commands, we have currently selected two possible
computers. The first one is the ISIS On Board Computer (iOBC) with a flight heritage starting in 2014.
The second one is the Cube Computer.

7. The iOBC runs with a 400 MHz 32-bit processor on a 3.3 V power supply. Its mass is 94 grams,
and has dimensions 96 by 90 by 12.4 mm. It is designed to either have two 8GB high reliability SD cards
or two standard SD cards of any size. Since the iOBC has heritage, its technology readiness level (TRL) is
roughly 6 to 7.
The Cube Computer runs with a 4-48 MHz 32-bit ARM Cortex-M3-based MCU on a 3.3 V
power supply. It has a 2 GB MicroSD socket for storage. It weighs 50 to 70 grams, and has dimensions 90
by 96 by 10 mm. The Cube Computer is newer so its TRL is 6 at best.
Thermal Management
Structure will be the biggest contribution to thermal management on the cubesat. Utilizing
a thicker structure around the payload will spread heat better. Because heat is transferred in space via
radiation, removing points of contact to prevent unwanted conduction will also be done. This can be done
utilizing tray inserts for payload. Materials will also be taken into account. Radiators placed over hot
components will be used to disperse heat to other components. The Sun’s thermal energy will also allow
heating of components during transit.
Conclusion
After multiple trade-off studies, the most feasible configuration of components and trajectories
has been determined. The instrumentation for the mission has been found to be weight and volume
efficient, a challenge considering that most cubesat technologies are not currently developed for Mars
operation. The greatest challenges in weight and volume were the impactor, propulsion, and power. These
rely on innovative but very possible technology. The primary science objective will be met with the
impactor/spectrometer combination. The secondary objective will be met with the onboard camera.
Communications will be relayed through MRO over the DSN.

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