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

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.

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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.

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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

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Mission Science Value Key science questions are

OriginCompositionRelationship 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

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Science Traceability Matrix

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Science ObjectivesMeasurement 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.

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

- The CubeSat must pass within 600 km of the plume ~1 hr after impact

- GNC, ADCS, propulsion requirements

- The CubeSat must analyze the 1.3 μm absorption and absorption levels of volatiles

- ADCS, spectrometer, C&DH requirements

- The CubeSat must transmit the data to the DSN

- Comms requirements

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OV-1

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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

- Observer burn ΔVo ~ 75.17 m/s

- Causes observer to arrive an hour after impact

- Flyby of observer

- Data collection

- Post mission objectives

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VIDEO HERE

Good window

Optimal case

Required ΔVc over one Deimos orbital period

Trajectory: Lining up with Deimos

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- 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

Spacecraft Architecture Overview

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- 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

- Propulsion: chemical

- Payload: spectrometer

- 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

- Payload:4 kg empty mass

6U CubeSat

ArchitectureOverview

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Overall Dimensions Impactor Dimensions Observer Dimensions

205.1x357.3x103.7 mm 205.1x153.7x103.7 mm 203.7x203.7x103.7 mm

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

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Impact Design

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

Flight Systems

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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

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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

- Custom battery protection circuitry

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

PropulsionObserver

- Aerojet Rocketdyne 2U MPS-130

- Chemical Monopropellant: AF-M315E

- Expected Isp of 240 seconds

- Green Propellant

- Available V = 229 m/s𝚫- 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 V = 39 m/s for corrections𝚫- Assuming Total Impactor Mass: 4.5 kg

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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)

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Telecommunications

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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

- X-band transpond

Pictured Above: Iris Transponder

Pictured Above: 4x4 Graphical representation of Tx patch.

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

- 2 GB MicroSD socket

- Redundant clocks

- Heritage from ADCS OBC on QB50 precursor satellites and DeorbitSail

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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

- Two 64 MB NOR flash

- IPC-A-610 Class C assembly certification

Payload Separation:NiChrome Wire Cutter

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- 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 Propulsion Cubesat Experiment (TEPCE)

- Total Cost per Unit: $166.21

System Engineering

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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

StructureAluminum 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

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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

EPSClyde 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

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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

Impactor Power Budget

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Impactor Power BudgetAverage Component Estimated Draw

SubsystemCBE Power (W)

Contingency (%)

MEV Power (W)

Structure and Mechanisms 0.00 0.20 0.00Thermal Control 0.00 0.20 0.00Power (inc. harness) 0.00 0.10 0.00On-Board Processing 0.55 0.05 0.585Attitude Determination and Control 2.00 0.15 2.30Propulsion 10.00 0.05 10.5Guidance 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

Telecom Link Budget, Data Volume and Return Strategy

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- 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.

Thermal Energy Balance and Management

Observer + Impactor Observer Impactor

α = absorbed 0.92 0.92 0.92

ε = emitted 0.85 0.85 0.79So = Earth Solar Flux 1370 1370 1370So = Mars Solar Flux 608.9 608.9 608.9A=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-8Watts (min) 25.69 25.69 .55Watts (max) 26 52 14.55Watts (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⋅ ⋅ ⋅

ConfigMax Tolerable Temperature (°C) Part

Min Tolerable Temperature (°C) Part

Observer + Impactor 40

Argus Spectrometer 5

Rocketdyne MPS-130

Observer 40Argus Spectrometer 5

Rocketdyne MPS-130

Impactor 40Clyde Space Battery -10

VACCO End-Mounted MiPS

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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

- Reduces radiation by 90% 3.15 rads

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Risk Identification & Mitigation1. 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

a. Plume is adjusted to be larger than needed by having a heavy 4 kg impactor.

5. Power failure

a. Contingency 12% for peak power requirements

6. Temperature failure

a. Spacecraft passively maintains correct temperature ranges

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Management, Schedule, Cost

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Program Schedule

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Cos

t Est

imat

e

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Total Project Cost

$3,783,955

With Contingency

$3,947,944

Cost: Personnel

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$601,955

$173,363

$669,794

Cost: Equipment

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17 18 19 20 21

Year of Purchase

Cost: Other Direct

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$100,000

$5,000

$401,877

$31,480

$12,949

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

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Questions?

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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].

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WBS Breakdown

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Cost Estimation

Observer Detail Power Budget

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Average Component Estimated Draw Maneuver Cruise Mode Science ModeDownlink 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 0Structure and Mechanisms 5.83 20.00 7.00 0 1 0 0On-Board Processing 0.13 5.00 0.14 1 1 1 1Attitude Determination and Control 0.50 15.00 0.58 1 1 1 1Propulsion 11.00 5.00 11.55 1 0 1 1Communications (Uplink) 12.00 15.00 13.80 1 1 1 0Communications (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

Link Analysis Detail

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Critical PathConcept 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

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