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SmallSat Mission Design School

Ball Aerospace

H1 Absorption in the Dark agES

Pioneers Mission Concept Presentation

Pre-phase A - July 30, 2021

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Cosmological models remain largely untested between the Dark Ages and Epoch of Reionization

HADES Main Goal

Provide key observational measurements for the foregrounds at radio frequencies in the range of 1-100MHz

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HADES: H1 Absorption in the Dark agES

Principal Investigator

Maryame El Moutamid

Senior Research Associate - CCAPS

Deputy Principal Investigator

Amit Vishwas

Research Associate – CCAPS

HADES lunar orbit provides efficient access to the radio zone above the lunar far side.

HADES is shielded from:

- Earth’s ionosphere

- Radio frequency interferences

- Solar interference.

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Radio Astronomy from the Moon

4

Current

Future

FARSIDE

Been of interest since 1960s!

Netherlands-China

Low Frequency Explorer

RAE 1/A (1968)

RAE 2/B (1973)

Past

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Radio Astronomy from the Moon

5

Current

Future

FARSIDE

Been of interest since 1968!

Netherlands-China

Low Frequency Explorer

MAY 1988

RAE 1/A (1968)

RAE 2/B (1973)

Past

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The Foreground Problem in Long Wavelength Radio Astronomy�

Cosmic baryon evolution during the Cosmic Dawn and Reionization results in spectral distortions of the redshifted 21-cm line of neutral hydrogen.

Appears as an absorption feature when observed against the CMB
Encodes within the nature and timing of when the first astrophysical sources came to be.

Detection requires precise methods to model the galactic and extragalactic radio sky, which is orders of magnitude brighter.

Ref: Pritchard & Loeb (2010)

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Why we need to go to the Moon

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Credit: Plice+2018 [DARE Mission Factsheet]

Man-made RFI is a large impediment to conduct sensitive Radio Astronomy observations.
Ionosphere distorts all low frequency observations and prevents high fidelity observations at <30MHz.
Far+Dark side of the Moon is likely the best location to conduct low frequency radio observations in the inner solar system.
Opening Unexplored Discovery Space – Seeing the Universe in a completely new light!

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Assessing the Lunar RF Environment

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Credit: Plice+2018 [DARE Mission Factsheet]

Prediction:

Infrastructure development plans already in place would potentially destroy this pristine environment
In order to execute this science case, we need rapidly deployable capabilities to support flagship observatories that are being planned for deployment in Lunar orbit or on the surface.

Bassett+2020

Takahashi 2002

Occultation observations done in the 70s & mid-90s showed how well can the moon block RFI (>60dB suppression)

Enable Data-driven pursuit of Radio Science from the Moon

Evidence:

EM Simulations predict Moon acts as a natural shield to an otherwise loud radio environment

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Science by HADES

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�

Project Scientist

Stella Koch Ocker

PhD Student - ASTRO

Dep. Project Scientist/Instrument

Ngoc Truong

PhD Student - EAS

Instrument Lead

Andrew Ridden-Harper

Research Associate - ASTRO

Data Processing

Nicholas Sitaras

Alexander Loane

MEng Student - MechE

Data Processing

Science Team	STM	Motivation	Sensitivity	Science Instrument	Data processing	Ancillary Science

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Science Goals	Science Objectives	Physical Parameters	Observables	Instrument Requirements	Projected Performance	Mission-Level Requirements
Characterize radio foregrounds from 1-100 MHz to determine feasibility of ground and space-based HI intensity mapping experiments probing the Dark Ages and Epoch of Reionization. Related NASA Objectives from the Astrophysics Roadmap and Astro2020 Decadal Survey: Probe structure of dark ages & epoch of reionization Characterize 1st starlight in the universe Test cosmological models	Determine whether HI absorption by spinning dust is a significant foreground for ground and space-based HI measurements.	Total, all-sky averaged brightness temperature measured by the antenna as a function of frequency between 1-100 MHz.	A ~50 mK absorption feature in the 20-80 MHz range produced by interstellar dust.	Frequency resolution: ≤ 50 kHz��Min. Frequency: 1 MHz��Max. Frequency [Baseline]: 100 MHz Max. Frequency [Threshold]: 90 MHz��System Temperature: <350 K��Sampling Time: <= 1 min.��Antenna Length: 3.3 meters��	Frequency resolution: 50 kHz��Min. Frequency: 1 MHz��Max. Frequency: 90 MHz��System Temperature: 250 to 330 K��Sampling Time: 1 min��Antenna Length: 3.3 meters	Total observation time ~560 hours on lunar far side from waning through waxing gibbous. Antenna boresight at Dec. of –70+/-10 deg & R.A. between \|20+/-10\| and \|80+/-10\| deg Location: +/- 65 degrees from the anti-Earth point to overlap with FARSIDE. Altitude <100 km to optimize lunar shielding from terrestrial RFI RFI/EMI <-80 dB suppression to avoid spacecraft noise.
	Measure the Galactic synchrotron spectrum from 1-100 MHz.		All-sky averaged brightness temperature of Galactic synchrotron radiation.
	Distinguish terrestrial and solar-origin radio frequency interference (RFI; e.g. peak frequency & bandwidth, temporal duration, & intensity).	Total RFI intensity as a function of time or orbital phase. RFI sources include terrestrial-based radio sources (e.g., satellite telecommunications) and solar radio bursts.	Frequency-time dynamic spectra covering 1-100 MHz with a sampling time <= 1 minute

Science Team	STM	Motivation	Sensitivity	Science Instrument	Data processing	Ancillary Science

Science Traceability Matrix

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Peering into the early universe

Chief observable = brightness temperature (K)

Total intensity referenced to a blackbody

Depth of HI absorption feature --> hydrogen spin temperature & ionization fraction
Standard cosmological models predict central frequency & brightness temperature

Departures from standard cosmology constrain exotic physics, e.g. dark matter interactions

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FARSIDE Concept Study Report (2019)

Signal	T	Ref. Frequency
Dark Ages Trough	-40 mK	1 - 40 MHz
Cosmic Dawn Trough	-200 mK	50 - 100 MHz

Science Team	STM	Motivation	Sensitivity	Science Instrument	Data processing	Ancillary Science

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The Foreground Problem

12

Measurements of the cosmological HI (21-cm) line require subtracting foregrounds that are up to ~6 orders of magnitude brighter.

Our scientific objectives target the following foregrounds:

Galactic synchrotron emission
Dust absorption
Terrestrial/solar radio frequency interference (RFI)

Science Team	STM	Motivation	Sensitivity	Science Instrument	Data processing	Ancillary Science

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Astrophysical Foregrounds: Galactic Synchrotron Radiation

13

Haslam et al. (1982)

Galactic synchrotron radiation = brightest foreground feature

All-sky averaged 1D spectrum

Pointing requirements driven by target threshold for Galactic foreground (~5000 K @ 50 MHz)

Science Team	STM	Motivation	Sensitivity	Science Instrument	Data processing	Ancillary Science

Thyagarajan et al. (2015)

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Draine & Miralda-Escude (2018)

Absorption by interstellar dust may mimic/broaden the cosmological HI line, but has not been searched for

Signal	T	Ref. Frequency
Dark Ages Trough	-40 mK	1 - 40 MHz
Cosmic Dawn Trough	-200 mK	50 - 100 MHz
Galactic Synchrotron	5000 K	50 MHz
Dust Absorption	-50 mK	80 MHz

Astrophysical Foregrounds: Interstellar Dust

Science Team	STM	Motivation	Sensitivity	Science Instrument	Data processing	Ancillary Science

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Solar & Terrestrial RFI

Solar radio bursts attenuated when Moon un-illuminated
Terrestrial RFI attenuated on lunar far side
Best attenuation <100 km above lunar surface
Observable = frequency-time dynamic spectra of RFI intensity w/ sampling time <= 1 min

measured both on far & near side of the moon

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Signal	T	Ref. Frequency
Dark Ages Trough	-40 mK	1 - 40 MHz
Cosmic Dawn Trough	-200 mK	50 - 100 MHz
Galactic Synchrotron	5000 K	50 MHz
Dust Absorption	-50 mK	80 MHz
Solar radio bursts	<10^(15) K	50 MHz

Bassett et al. (2020)

Science Team	STM	Motivation	Sensitivity	Science Instrument	Data processing	Ancillary Science

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Sensitivity

16

Radiometer equation:

RMS noise ~5 mK/(0.5 MHz) --> 560 hours on lunar far+dark side

100 hours total for calibration before/after each science observation

calibration

noise from spacecraft components

T(sky)~5000 K

relative suppression threshold: -80 dB

T(receiver)~300 K

Science Team	STM	Motivation	Sensitivity	Science Instrument	Data processing	Ancillary Science

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Sky Signal

Architecture based on Parker Solar Probe heritage with lower mass by using more digital components

JIB Monopole by Northrop Grumman (first introduction in 1963, more than 1000 units deployed on orbit)

Instrument

Science Team	STM	Motivation	Sensitivity	Science Instrument	Data processing	Ancillary Science

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Low frequency filter

Analog to Digital converter (ADC)

Field Programmable Gate Arrays (FPGAs)

Dipole antenna

Length: 3m

Resonant Freq. ~50 MHz

Filter out < 1 MHz. Lunar ionosphere plasma frequency 0.2 - 1 MHz

Sampling rate ~200 Msps

Time average, channelize data, create frequency spectrum

Output

Receiver digital backend

Characterize receiver with regular calibration
Digital backend can utilize COTS Integrated On-board Computing System. E.g., Xilinx Zynq ARM Cortex
Receiver outputs handled by Data Processing

Calibration unit:

Input known signal

switch

Low-noise amplifier (LNA)

Relative receiver noise is reduced

Instrument

Science Team	STM	Motivation	Sensitivity	Science Instrument	Data processing	Ancillary Science

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Mission Minimum Science Criteria = 1.6 MB/Day (within proper lunar phase)

Data Processing

Raw Science Data

Frequency Spectrum Data

Data Storage paired with timing data

Analysis of data by science team on ground

Distribution and Publication

Science Team	STM	Motivation	Sensitivity	Science Instrument	Data processing	Ancillary Science

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Ancillary Science

Data products useful for broad community stakeholders

20

Jupiter's aurora taken by Hubble. Credit: nasa.gov

Jupiter's aurora

synchrotron spectrum

Galactic magnetic fields & energy transport in ISM

dust absorption

joint observations w/ LOFAR, post-Juno

nanoparticle abundance in ISM

Science Team	STM	Motivation	Sensitivity	Science Instrument	Data processing	Ancillary Science

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Management

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Program Manager

Trevor Foote

PhD Student, Astro

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(Emily Matteson)

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Budget

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The reserves are based on assessed cost risks:

For Medium cost risk, we allocated 50% reserves
For Low cost risk, we allocate 30%
For Very Low cost risk, we allocate 15% or less

Total costs for Phase A are $533k, which include formation of CSR, and subsequent three month pause

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Schedule

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Pre-Phase A: Concept Studies

Phase A: Concept and Technology Development (CSR - Concept Study Report)

Phase B: Preliminary Design and Technology Completion (PDR – Preliminary Design Review)

Phase C: Final Design and Fabrication (CDR – Critical Design Review; SIR – System Integration Review)

Phase D: System Assembly, Integration and Test, Launch

Phase E: Operations and Sustainment

Phase F: Closeout

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Technical

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

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Concept of Operations

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GTO Rideshare

Attitude Acq. and Panel Deployment

24 months

GTO to Lunar Orbit

5 months

Threshold/Baseline Science phased with Coms

Science Orbit

< 2 months

Disposal

Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Systems Engineers

Lead: Katherine Wilson

PhD Candidate, AeroE

Deputy: Emily Matteson

Junior Undergrad, MechE

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Trajectory Subsystem: GTO Orbit

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

A. GTO

B. GTO to Lunar Transfer

C. Lunar Capture

D. Science Orbit

E. Disposal

Trajectory Engineers

Lead: Kalani Danas Rivera

PhD Candidate, AeroE

Deputy: Grace Genszler

PhD Student, AeroE

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Trajectory Subsystem: GTO Orbit

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Launch Date:

August 2024

Ride Share

Services with 12U capabilities: Tyvak and Spaceflight
Possible vehicles: ULA Vulcan and SpaceX Falcon 9

Launch Vehicle

Elise Eckman

PhD Student, AeroE

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Trajectory Subsystem: GTO to Lunar Transfer

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Continuous Low Thrust from GTO to Moon Sphere of Influence
Delta V = 3.5 km/s
Transfer Time 600 days
Earth Eclipse

~100 Hours Total Eclipse Time
~2 Hours Maximum Single Eclipse Time

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Trajectory Subsystem: Lunar Orbit Insertion

SMART-1 Mission Case Study

Lunar Capture

4 Lunar Flybys

Lunar Descent into Science Orbit

Continuous Low Thrust Burn
For HADES transfer time 135 days

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

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Trajectory Subsystem: Science Orbit

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Total Orbital Period

120 min.

~60 min.

Data Downlink

Solar Power Generation

~48 min.

Science Collection

Nearside

Farside

Circular Equatorial

Orbit

At 60 km Altitude

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Trajectory Subsystem: Station Keeping

Science orbit requirements

20-100 km altitude
+/- 65° inclination

Station keeping orbit design

28.5-91.5 km altitude
0°-2° inclination
0-.035 eccentricity

No station keeping maneuvers required
Duration = 5 months

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

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Trajectory Subsystem: Disposal

Intentional crash land onto the surface of the moon

Low delta V usage
Allow for data downlinking during deorbiting

Continuous low thrust trajectory
Duration = 3.7 days
Delta V = 20 m/s
Target site selection

Comply with planetary protection guidelines
Chang'e-1 (1.5°S 52.36°E)

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

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Propulsion and Launch Subsystem

Self-Contained Mission from GTO to Lunar Orbit

Novel: Propulsion closes the loop for first ever GTO to lunar transfer for a small satellite.
Will set new delta V record for small satellites.

Challenging Propulsion Requirements GTO to Lunar Orbit

Δv with contingency = 4.35 km/s .
Max Power with contingency: 137 W.
Max Volume: 5U.

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Propulsion Engineers

Lead: Carl Geiger

PhD Student, AeroE

Deputy: Elise Eckman

PhD Student, AeroE

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Propulsion and Launch Subsystem

Requirements:

Power ≤ 137 Watts

Volume ≤ 5 U

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

System	Enpulsion Micro R3 (x1)	Enpulsion Micro R3 (x2)	BGT-X5	Halo Micro Hall Effect Thruster	BHT-200
Transfer Time (days)	1332	735	75	224	57
Total system mass (kg)	3.9	7.7	21.6	15	5.9
Power Required (W)	70	137	80	180	200
Volume Required (U)	2.25	4.5	12	6.5	6.5
Propulsion type	FEEP	FEEP	Monopropellant	Hall Thruster	Hall Thruster
TRL	7	7	7	6	6

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Propulsion and Launch Subsystem

Enpulsion Micro R3 (x2)

Flight proven as of March 15th, 2021
Ongoing lifetime test at >24,000 hours
Minimal envelope and power requirements
0.635 mN of thrust each

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https://www.enpulsion.com/wp-content/uploads/ENP2018-002.H-MICRO-Thruster-Product-Overview.pdf

Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

T

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Propulsion and Launch Subsystem

VACCO CuSP MiSP

Very small envelope (0.3 U)
Relatively high total impulse (69.4 N-s)
Completely inclusive (propellant tank, microprocessor, etc.)
4 independently operatable thrusters for angular

momentum dumping along all three axis

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https://cubesat-propulsion.com/cusp-propulsion-system/

Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

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Propulsion and Launch Subsystem

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Maneuver	Thruster	Total Impulse (N/s)	Delta V (m/s)	Time (days)
GTO to Science Orbit	FEEP	80,800	4,350	735
*Attitude Momentum Dumping	Cold Gas	24.5	1.36	N/A
Science to Deorbit	FEEP	360	20	3.7

Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Summary

*Cold gas requirements for pointing during GTO to science orbit not considered at this time, will be left to future work

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ADCS Subsystem

Initial Detumbling: < 2 hr
Pointing requirement: 0.715° (40% safety margin)

ADCS Design (TRL 9)

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

ADCS Engineers

Lead: Aneesh Heintz

PhD Candidate, AeroE

Deputy: Andy Tan

Undergrad, MechE

Science	Trajectory Control	Communications	Power
6°	1°	20°	18.5°

Actuators	Sensors	Att. Det. (EKF)	Orbit Det. (EKF)
Non-singular 3-axis active control 4 Reaction Wheel Assembly	Star tracker 3 gyros Sun / horizon sensor	Near Side: All sensors Far side: Gyros & star tracker	Near side: Two-way Doppler Ranging using DSN & maybe NEN for support (upgrades expected as of 2017) Far side: Propagate EKF

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ADCS Subsystem

Disturbance Torque Analysis

Max est. disturbance torque:

Solar Pressure, Gravity Gradient, etc.
2.8×10^-7 Nm

Required momentum storage:

Accumulation over ¼ orbit
6.6×10^-4 Nms

Pointing Maneuver Analysis

Max Slew torque:

2.5×10^-3 Nm

Max Overshoot:

5% of total maneuver

Max Settling time:

10-80-10 slew profile
182 s

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Uncertainty Analysis

Reaction Wheel Control Accuracy:

Best case: 0.002°

Sensor Estimation Accuracy:

Worst case: 0.2°
Best case: 1.6×10^-3° (6")

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Thermal Subsystem: Problem

Problem: Maintain the temperature of all components of HADES within operational limits.

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Environment Protection Engineers

Lead: Fernanda Fontenele

PhD Candidate, MechE

Deputy: Emily Matteson

Junior Undergrad, MechE

PLICE, Laura; GALAL, Ken, et al.

Sun Exposure in Lunar orbit

Plan of action:

Determine cycles of temperature of satellite in Lunar orbit
Select thermal control: passive or active? Where to place them?

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Thermal Subsystem: Transient Analysis

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1- Model: create a Lumped Thermal Capacity model to represent the satellite.

3 - Simulations: verify Model 1. Determine thermal parameters accurately.

2- Results: get temperature cycles. Select thermal control.

Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Temperature

Time

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Thermal Subsystem: Selected design

Current Design

Temperature of satellite: from -60°C to 70°C (body -5°C to 35°C).
Passive thermal control: Alodine coating (increase of body min temp to 9°C)
Active thermal control: built in heater on battery + 1 extra safety heater (autonomous control)
Temperature measurement: 1 thermistor per subsystem + redundant thermistor

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Temperature of satellite with Al coating

Temperature cycles of satellite with no thermal control

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Communications and Ground Subsystem: Overview

Downlink Science data

Use Near Earth Network (NEN)
Transmit when on near side of orbit (30 min duration)
Transmit total of 1.6MB of data per day

Communicate navigation data

Use Deep Space Network (DSN)
Transmit when on near side of orbit (5 times per orbit, ~1 min duration each)
Rate of 128 kbps

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Coms and Ground Systems Engineers

Lead: Jordan Sandell

Undergraduate, MechE

Deputy: Grace Genszler

PhD Student, AeroE

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Communications and Ground Subsystem: �Selected Design

Transceiver: JPL/Iris Transponder

Dimensions: 10.5 x 10.1 x 5.6 cm (~0.5U volume)
Mass: 1.2 kg
Frequency Band: X (8000-8500 MHz)
Power Consumption: 12.6W/35W (receive only/full transponder mode)
Power Output: 4W
Modulation: BPSK
TRL: 9

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

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Communications and Ground Subsystem: �Selected Design

Antenna: AntDevCorp Patch Array

Dimensions: 4.70 x 4.70 x 1.40 cm
Mass: 300g
Frequency Band: X (8000-8500 MHz)
Gain: 9
Pointing requirement: 20º
TRL: 9

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

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Communications and Ground Subsystem: �GTO to Lunar Orbit Coverage

Full Coverage

DSN Locations

California
Spain
East Australia

NEN Locations

Hawaii
West Australia
Virginia
South Africa

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

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Communications and Ground Subsystem: �Station Keeping and Disposal Coverage

Minimal No Coverage Zone Crossing

Change downlinking schedule

DSN Locations

California
Spain
East Australia

NEN Locations

Hawaii
West Australia
Virginia
South Africa

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

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Communications and Ground Subsystem: �Link Budget

Location: Dongara, Western Australia

Antenna Diameter: 13 m

Effective Data Rate: 10,000 bps (8 kbps)

Eb/N0: 9.74 dB

Margin: 10.95 dB

Bit Error Rate: 10^-7

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

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Power Subsystem

Power Budget (GTO-Lunar)

180 W Max Solar Power

150 W Average Consumption

20% Contingency

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Power Engineers

Lead: Liam Webster

Senior Undergrad, MechE

Deputy: Dohun Kim

Senior Undergrad, MechE

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Power Subsystem

Power Budget (Lunar Orbit)

30.78 W Orbit Average Consumption

Transponder is the largest consumer

40% Contingency

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

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Power Subsystem

NanoPower TSP

12 Total Panels

6 Per Array

2 DOF Sun Tracking

Battery

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PSU

Lithium-Ion Battery: NanoPower BP8

86 Wh
Avg. Depth of Discharge 20%

Power Supply Unit: NanoPower P80

Handle solar generation up to 300 W

1980mm

208mm

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Power Subsystem

EMI Shielding: Instrument Requirements

Spacecraft RF emissions should be suppressed by at least 80 dB
Frequency Range: 1MHz to 200MHz (according to ERD)

54

DARE's approach:

"Two-step lid using RF gaskets … components housed inside an additional faraday cage"

	Copper Foil	Aluminum Polyester Foil	Nickle Conductive Spray Paint	BeCu Gaskets + Al Faraday Cage
dB Range	65	50 – 63 (63 – 70)	35 – 65 (65 – 90)	102 – 130
Frequency Range	300 kHz – 2.5 GHz	30 MHz – 200 MHz (200 MHz – 1500 MHz)	1 MHz – 200 MHz (0.01 MHz – 1 MHz)	1 MHz – 400 MHz
Cost ($)	10.99	2.11	21.95	20 (+ 150)

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Structures and I&T Subsystem

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Structures and I&T Engineers

Lead: Liam Alexis

PhD Student, AeroE

Deputy: Allie Voltaggio

Junior Undergrad, MechE

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

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Structures and I&T Subsystem

Dark Side Operation

Science Phase
Primarily on the dark side of the moon

Solar Operation

Battery Charging
Downlink Phase

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Science Phase
Primarily on the dark side of the moon

Battery Charging
Downlink Phase

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Structures and I&T Subsystem

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Testing of qualified components

Qualification of proto-flight hardware

NASA General Environmental Verification Standards (GEVS)

GSFC-STD-7000B

Launch provider requirements

OR

Goddard Technical Handbook GSFC-HDBK-8007 – Mission Success Handbook for Cubesat Missions

First four thermal cycles
Random vibe
First 500 hours of operation
Close-out inspection
Early holistic risk assessment
“iphone” photography
informal independent SME review (graybeard mentoring)
spare printed circuit board” for coupon for future DPA

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Command & Data Handling Subsystem

Redundant Flight Hardware
Rad-hardened external clock
Integrated ADCS available for COTS parts

SatBus 3C2

Storage capabilities up to 2 GB for shorter integration times

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

CDH Engineer

Alexander Loane

MEng Student- MechE

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Review: Feedback and Questions

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Supplementary Material

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Mission Risks

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Risk	Mitigation Strategy	Approach
Failure to survive environment from GTO to EOL Failure to deploy instrument antenna Failure to deploy solar panels Loss of communication	Use space-rated hardware when possible with redundancy and/or heritage parts used for similar duration if necessary Deployment testing, selection of deployment mechanism with flight heritage Deployment testing and placement of ADCS sensors and instrument away from the panels Communication system testing, data downlinking each science orbit, health checks, data storage capacity	Mitigate Mitigate, Watch Mitigate, Watch Mitigate

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Environmental Protection

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

Source of Risk	Approach
Solar radiation/cycles, Van Allen belts, Galactic cosmic rays Outgassing Discharge	Radiation-hardened electronics selected when possible, redundancy otherwise Time for "bake out" has been integrated, flight heritage requirement to avoid risk from materials Selection of conductive coatings in Phase A for grouped electronics will facilitate charge distribution

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Communications and Ground Systems

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

South Point, Hawaii

Wallops Island, Virginia

Hartebeesthoek, South Africa

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Communications and Ground Systems

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Mission Overview	1: Trajectory	2: LV and Propulsion	3: ADCS	4: Environment/ Thermal	5: COMS/ Ground	6: Power	7: INT/ Test	8: CDH

DSN Uplink

DSN Downlink

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Conclusions

Relevance to NASA

HADES realizes NASA’s strategic objectives to “... explore how (the Universe) began and evolved ...”

HADES executes small-scale mission described in Astrophysics Roadmap: “Mapping the Universe’s hydrogen clouds using 21-cm radio wavelengths via lunar orbiter from the far side of Moon.”

Novel Small Sat trajectory from GTO to Lunar Orbit

Pathfinder to characterize Lunar Radio Frequency Environment

Study feasibility of conducting observations of the Early Universe

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SmallSat Mission Design School

Mentors:

Ball Aerospace

HADES MISSION

Pioneers Mission Concept Review Presentation