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

1. ADIOS
-
A Deimos Impact & Observation Spacecraft
Team 3
Jeff Anderson, Thomas Blachman, Andrew Fallon, John Franklin, Samuel Gaultney,
David Habashy, Brian Hardie, Brandon Hing, Zujia Huang, Sung Kim, Jonathan Saenger

2. Mission Goal
Primary: Direct an impactor into Deimos at high velocities to launch a plume of
surface and subsurface debris into space. The released plume will be analyzed by
a passive infrared spectrometer to determine the composition of Deimos. This will
determine whether Deimos is a C or D type asteroid, or Mars ejecta.
Secondary: Prebiotic volatile concentrations will be analyzed to determine the
potential asteroid contributions to early life.
Alternative: Close Proximity Imaging of one face of Deimos with passive
spectrometry of surface composition or total satellite impact with spectrometry
conducted by Mars satellites.
2

3. 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.
3

4. Key Mission Requirements
- Shall be ready for launch by July 14th, 2020
- Shall not exceed $5.6 M in total cost
- Shall not exceed 14 kg for all components
- Be able to deliver the impactor to the surface of Deimos 50 minutes before the observer
- Be able to deliver the impactor to Deimos at a speed no less than 3.5 km/s and a mass
of 4 kg to produce a sufficient plume size of 0.25 km x 0.25 km
- Be able to determine the 1.3 µm absorption levels of the plume as well as the
absorption levels of volatiles
- Be able to point the spectrometer at the plume for a minimum of 30 seconds at a range
of no more than 600km
- Be able to relay all spectrometer data back to Earth via the DSN
4

5. Mission Science Value
Key science questions are
Origin
Composition
Relationship to other solar system materials.
Are the moons possibly re-accreted Mars ejecta [or] primitive, D-type bodies? Spectrometry can answer this question.
“Resolving the debate concerning the compositions (and likely origins) of... Deimos may be relevant to understanding the
early history of Mars...if they turn out to be related to volatile-rich asteroids...they may be the surviving representatives of a
family of bodies that originated in the outer asteroid belt or further, and reached the inner solar system to deliver volatiles
and organics to the accreting terrestrial planets.”
-Decadal Survey
5

6. Science Traceability Matrix
6
Science Objectives
Measurement
objectives
Measurement
Requirements
Instrument
Requirements Instruments Data Products
Deimos
Internal composition Measure ratio of iron in
internal composition
Spectronomy
measurements for 160
seconds
Be able to measure the
1.3 µm absorption
levels of the plume
ARGUS Spectrometer Graphs of Spectronomy
Readings
Internal volatiles Determined the amount
and type of subsurface
volatiles
Spectronomy
measurements for 160
seconds
Be able to measure the
1.0 µm - 1.63 µm.
absorption levels of the
plume
ARGUS Spectrometer Graphs of Spectronomy
Readings
Decadal Survey: “Are the moons possibly re-accreted Mars ejecta? Or are they possibly related to primitive, D-type bodies? These
questions can be investigated….mission that includes measurements of bulk properties and internal structure.”
MEPAG goals Investigation A3.1: “Characterize organic chemistry, including (where possible) stable isotopic composition and
stereochemical configuration. Characterize co-occurring concentrations of possible bioessential elements.”
Mission Objective: Measure the internal subsurface composition of Deimos to determine its origins and organic volatile levels.

7. Requirement Flowdown
- Project ADIOS will determine the surface and subsurface composition of Deimos through
spectrometry using a CubeSat and detachable impactor
- The impactor shall strike Deimos with a mass and velocity sufficient to generate an analyzable
plume
- The impactor must detach safely from the CubeSat
- Separation mechanism requirements
- The impactor must navigate to Deimos
- GNC, ADCS, propulsion requirements
- The impactor must arrive with a mass of 4 kg and a speed of 3.5 km/s
- The CubeSat shall perform spectrometry on the generated plume and transmit the data back
to Earth for analysis
7

8. OV-1
8

9. Trajectory：
Overview and Maneuvers - Separation from Mars 2020
- Initial burn ΔVi ~ 41.46 m/s
- Occurs after 4 days
- Achieve Martian altitude of 30,000 km
- Achieve inclination of 0° relative to
Deimos’ orbit
- Impact burn ΔVc ~ 19 m/s
- at Mars’ SOI
- Achieve impact with Deimos
- Separation of Observer and Impactor 9
VIDEO HERE

10. Good window
Optimal case
Required ΔVc over one Deimos orbital period
Trajectory:
Lining up with Deimos
10
- Retrograde Hyperbolic Trajectory for
maximum impact velocity
- Over 12 hours window available each 30
hours (Deimos’ orbital period) to keep ΔVc
low
- Adjustment to delay/advance arrival time
can be done at initial separation
Worst case
Optimal case
Satisfactory
Deimos

11. Spacecraft Architecture Overview
11
- 4U Observer Module
- Self-contained, self-controlled
- ADCS: star trackers, sun
sensors, reaction wheels
- GNC: DDOR
- Comms: transceiver
- C&DH: Cube Computer
- EPS: solar panels, batteries
- 2U Impactor Module
- Self-contained, self-controlled
- ADCS: star trackers, sun
sensors, reaction wheels
- GNC: camera
- C&DH: NanoMind A 3200
- EPS: batteries
- Propulsion: cold gas
6U CubeSat

12. Architecture
Overview
12
Overall Dimensions Impactor Dimensions Observer Dimensions
205.1x357.3x103.7 mm 205.1x153.7x103.7 mm 203.7x203.7x103.7 mm

13. Payload: Spectrometer
Selected Instrument: ARGUS
- Passive infrared spectrometer
- Operates in 1 μm to 1.7 μm range
- Extended range version goes to 2400
nm
- Range: 600 km
- FOV: 0.15°
- Power: 1.4 W
- Volume: 0.18U
- Integration Time Ranges: 500 μs to ~4
seconds
- Data transmitted in 100 ms
- Can adjust number of scans for co-
adding spectra
Requirements Necessary:
- Must have a spectronomy range of 1.0
µm to 1.63 µm.
- Physical range of greater than 400 km
- Size must be less than 2U
- Must make measurements in under 80
seconds
13

14. Impact Design
14
- Average Density of plume at arrival 0.02 kg/m3

15. Flight Systems
15

16. Structure
- Custom-built aluminum frames
- Insulating layers for thermal
containment
- Observer has 0.5U modules
attached to the central
propulsion frame
- Impactor has a single frame
- Components slot in individually
- Protection from 35 rads is
accommodated by 0.8 mm
aluminum on necessary parts
16

17. Power
Observer
- Clyde Space Deployable, Double-Sided
Solar Cells
- 5 mm Profile fits to 4U structure
- 40 W Peak Power at Mars, 20.8 W
Average Orbit Power
- Clyde Space FlexU CubeSat EPS
- Up to 12 Solar Panels
- 98% Efficient at 5 V and 3.3 V
Regulators
- Clyde Space 60 Wh Battery
- 10.4 Ah at 8.0 V to 6.4 V
Impactor
- Clyde Space FlexU CubeSat EPS
- Up to 12 Solar Panels
- 98% Efficient at 5 V and 3.3 V Regulators
- 3x Clyde Space 40 Wh Battery
- 10.4 Ah at 8.0 V to 6.4 V
- Custom battery protection circuitry
17Observer Solar Panel Configuration

18. Propulsion
Observer
- Aerojet Rocketdyne 2U MPS-130
- Chemical Monopropellant: AF-M315E
- Expected Isp of 240 seconds
- Green Propellant
- Available 𝚫V = 229 m/s
- Assuming Total Spacecraft Mass: 14 kg
- Cost Savings
- Simplified range operations
- Reduction of thermal management
Impactor
- VACCO End-Mounted 0.5U MiPS
- Cold-Gas Propellant: R134a
- Isp of 40 seconds
- Non-Toxic
- Available 𝚫V = 39 m/s for corrections
- Assuming Total Impactor Mass: 4.5 kg
18

19. ADCS
- BCT XACT
- 0.5 U
- 3-axis control
- Contains Star Trackers, Reaction
Wheels
- 1-sigma cross-axis pointing error
better than 8 arcseconds
- Pointing Accuracy: 0.003° (2 axis),
0.007° (3rd axis)
- Slew Rate: 10 deg/s
GNC
Observer
- Delta-DOR
- Utilize DSN and IRIS Comm.
System on CubeSat
- Used by ESA for interplanetary
missions such as Mars Express
Impactor
- MSSS ECAM-M50 (Camera)
19

20. Telecommunications
20
Iris V2
- Antenna
- 8x8 Tx Patch
- 1000-62 bps
- Capable of transmitting 5.16 MB
in less than 10 minutes
- Covers 2x2 U surface
- Rx patch integrated into TX board
- 1.2 kg, 0.5U
- 26 W at full transpond
Pictured Above: Iris Transponder
Pictured Above: 4x4 Graphical
representation of Tx patch.

21. Command and Data Handling
- Cube Computer
- Off-the-shelf
- Operating Voltage: 3.3V
- PC/104 Form Factor compatible with
CubeSat
- Internal and external watchdog
- 400 MHz processor
- Two 1 MB SRAM for data storage 21
Observer Impactor
- NanoMind A 3200
- Off-the-shelf
- Real Time Clock
- Operating Voltage: 3.3V
- 3-Axis gyroscope
- On-board temperature sensors
- 32 MB SDRAM
- 512 KB built-in flash

22. Payload Separation:
NiChrome Wire Cutter
22
- NiChrome Wire Cutter Release
Mechanism
- Created by Adam Thurn
- The two saddles (see green in model)
are only non-commercial parts
- Dimensions: 32 x 16.5 x 11.5 mm
- Average Vacuum Cut Time of Vectran
- 200 Denier: 2.6 Seconds
- 400 Denier: 6.2 Seconds
- Used on Tether Electrodynamics

23. System Engineering
23

24. Observer Mass Budget & TRLs
Subsystem Component (Quantity)
Current Best
Estimate (kg)
TRL
Contingency
(%)
Maximum Expected
Value (kg)
ADCS BCT XACT 0.91 9 5 0.956
Communication Iris V2 1.2 5 25 1.5
C&DH Cube Computer 0.07 9 5 0.074
EPS
Clyde Space FlexU EPS 0.148 8 10 0.163
Clyde Space 60 Wh Battery 0.475 8 10 0.523
Clyde Space 2U Deployable Array (4) 0.8 8 10 0.88
Payload Argus 1000 IR Spectrometer 0.23 9 5 0.242
Propulsion (Wet) Rocketdyne MPS-130 3.5 6 25 4.375
Structure
Aluminum Frame (2) 0.201 9 5 0.211
Fasteners (50) 0.25 9 5 0.263
Radiation Shielding .25 9 0 .25
Misc. Cables, Wires (20) 0.1 9 5 0.105
Subtotal (Dry) 6.834 8.239
Subtotal (Wet) 8.234 9.539 24

25. Impactor Mass Budget & TRLs
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 6 25 0.018
EPS
Clyde Space FlexU EPS 0.148 8 10 0.163
Clyde Space 40Wh Battery (3) 0.954 8 10 1.05
GNC MSSS ECAM-M50 0.256 7 20 0.307
Propulsion (Wet) VACCO End-Mounted MiPS 0.924 6 30 1.201
Structure
Aluminum Frame 0.617 9 5 0.648
Fasteners (25) 0.125 9 5 0.131
Radiation Shielding 0.15 0.15
Misc. Cables, Wires (10) 0.05 9 5 0.053
Subtotal (Dry) 3.725 4.252
Subtotal (Wet) 4.148 4.675
Maximum Expected Total Dry Mass (kg) 12.491
Maximum Expected Total Wet Mass (kg) 14.214 25

26. 26
Observer Power Budget
- Solar panels will provide
enough power for majority
of modes
- Battery will be fully charged
from Earth and will be used
during Downlink Mode
27 26.06
60
21.51 25.69
52.51

27. Impactor Power Budget
27
Impactor Power Budget
Average Component Estimated Draw
Subsystem
CBE Power
(W)
Contingency
(%)
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 0.585
Attitude Determination and
Control 2.00 0.15 2.30
Propulsion 10.00 0.05 10.5
Guidance and Navigation
Control 2.00 0.15 2.3
Total Power 14.55 15.68
- Only one Mode
- 120 Wh battery will allow for
multiple maneuvers since
propulsion will only use
power for minutes at a time
- Battery will be fully charged
from Earth

28. Telecom Link Budget, Data Volume and Return
Strategy
28
- Utilize 8x8 Tx Patch
- Opposition: 1000 bps
- Conjunction: 62 bps
- Total Data Accumulated:
- 5.16 MB
- Entire end of life utilized to
transmit data
- At peak rate, ~10 minutes.

29. Thermal Energy Balance and Management
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 .55
Watts (max) 26 52 14.55
Watts (heater) 0 10 0
Earth cruise 37.65199138
Mars cruise 0.4701006878 11.49890 -8.67
Mars full power 0.8267963944 23.38434 16.45723844
Q e = ε⋅σ⋅ Ar⋅Tr^4
Qa = So⋅α⋅A⋅cos(θ)+Watts+heater
Config
Max Tolerable
Temperature (°C) Part
Min Tolerable
Temperature (°C) Part
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
29

30. Radiation Shielding
- ADIOS will experience
approximately 35 rads during its
mission
- Calculated from Curiosity
measurements
- An adequate amount of aluminum
shielding will be applied to protect
vital components
- 0.8 mm thick
- 400 g 30

31. Risk Identification & Mitigation
1. Damage to key systems from Radiation
a. All components have radiation hardening for mission time
or are otherwise insulated.
2. Trajectory Mishap
a. 33% extra fuel for course corrections
b. Communication directly back to earth possible
3. Impactor Fails Separation
a. Surface Spectrometry
b. Redundant release system
4. Plume Size Failure
31

32. Management, Schedule, Cost
32

33. 33

34. Program Schedule
34

35. CostEstimate
35
Total Project Cost
$3,783,955
With Contingency
$3,947,944

36. Cost: Personnel
36
$601,955
$173,363
$669,794

37. Cost: Equipment
37
17 18 19 20 21
Year of Purchase

38. Cost: Other Direct
38
$100,000
$5,000
$401,877
$31,480
$12,949

39. Descope Options
- Use MRO or future spacecraft to do spectronomy
- Saves $49,000 for Argus and no longer need separate impactor
- Have impactor be unguided
- Saves $200,000 in component costs and reduces complexity
- Increases risk of missing.
- Reduce the amount of employees
- Cutting 2 graduate students saves $484,452.77 over 5 years
- Only do spectronomy of Deimos Surface
- Backup in case of impactor failure
39

40. Questions?
40

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50“NASA Space Flight Program and Project Management Handbook,” NASA, Hampton. (2014). https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20150000400.pdf. [Retrieved 5 November
2016].
51“On Board Computer,” https://www.cubesatshop.com/product/on-board-computer/. [Retrieved 28 October 2016].
52"Products- CS High Power Bundle" Clyde Space. N.p., n.d. https://www.clyde.space/products/47-cs-high-power-bundle-c-eps-80whrbattery. [Retrieved 14 October 2016].
53"Products- Double Deployed Solar Array" Clyde Space. N.p., n.d. https://www.clyde.space/products/27-2u-doubledeployed-solar-array. [Retrieved 14 October 2016].
54“Reaction Control Propulsion Module,” CubeSat Propulsion Systems, URL: http://www.cubesat-propulsion.com/wp-content/uploads/2015/10/reaction-control-propulsion-module.pdf. [Retrieved 1
November 2016].
55“Thoth Technology, Inc. ‘Argus 1000 IR Spectrometer Owner’s Manual,’” http://thothx.com/manuals/Argus%20Owner's%20Manual,%20Thoth%20Technology,%20Oct%2010,%20rel%201_03.pdf.
[Retrieved 25 September 2016].
56“Vision and Voyages for Planetary Science in the Decade 2013-2022,” Washington, D.C.: National Academies Press, ©2011. <http://solarsystem.nasa.gov/docs/Vision_and_Voyages-FINAL.pdf>.
[Retrieved 15 October 2016].
45

46. 46

47. WBS Breakdown
47

48. 48

49. 49
Cost Estimation

50. Observer Detail Power Budget
50
Average Component Estimated Draw Maneuver Cruise Mode Science Mode
Downlink
Mode
Subsystem CBE Power (W) Cont. %
MEV Power
(W) Duty Cycle Duty Cycle Duty Cycle Duty Cycle
Spectrometer 1.24 15.00 1.43 0 0 1 0
Structure and
Mechanisms 5.83 20.00 7.00 0 1 0 0
On-Board
Processing 0.13 5.00 0.14 1 1 1 1
Attitude
Determination and
Control 0.50 15.00 0.58 1 1 1 1
Propulsion 11.00 5.00 11.55 1 0 1 1
Communications
(Uplink) 12.00 15.00 13.80 1 1 1 0
Communications
(Downlink) 35.00 15.00 40.25 0 0 0 1
Total Power 52.46 59.51 26.06 21.51 25.69 52.51

51. Link Analysis Detail
51

52. Critical Path
Concept Studies: Jan. 2017 - Feb. 2017
Concept/Technology Development:
Mar. 2017-July 2017
Prelim. Design: Aug. 2017 - Mar. 2018
Final Design/Fabrication: Apr. 2018 - July
2019
Sys. AI&T: July 2019-July 2020
Launch & Ops: July 2021 - Mar. 2021
Decommissioning: Apr. 2021 - June 2021
52