SAGAN Mission to Study Atmospheric Gradients and Neutral forcing in Near-Earth Plasma

Study of Atmospheric Gradients and
Neutral forcing (SAGAN) Mission
Vaibhav Kumar
Tanish Himani
Swapnil Pujari
Mark Mote
Matthew Owczarski

Motivation/
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Mission
Overview
Flight System
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Management
Summary
Mission Relevance
• IMCC Request for Proposals calls for a mission
designed to “understand how the ionosphere is
driven by, and participates in, the global circulation of
plasma and energy throughout the coupled
ionosphere-magnetosphere system”
• Our Science Goal: To understand how lower
atmospheric wave energy, neutral forcing and current
drifts in the low altitude ionosphere affect the near-
Earth plasma.
2

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Mission Relevance
3
Space Studies Board,. The 2013-2022 Decadal Survey In Solar And Space Physics.
Division on Engineering & Physical Sciences: N.p., 2012. Web. 13 Sept. 2015. Solar And
Space Physics: A Science For A Technological Society.

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4
Space Studies Board,. The 2013-2022 Decadal Survey In Solar And Space Physics.
Division on Engineering & Physical Sciences: N.p., 2012. Web. 13 Sept. 2015. Solar And
Space Physics: A Science For A Technological Society.

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Mission Relevance
• The Decadal Survey committee recommends the
following:
– Science Goals: Determine the dynamics and coupling of
Earth’s magnetosphere, ionosphere and atmosphere and
their response to solar and terrestrial inputs.
– Guiding Principles: To make transformational scientific
progress, the Sun, Earth, and heliosphere must be studied as
a coupled system.
5

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NASA,. NASA 2014 Science Plan. NASA, 2014. Web. 13 Sept. 2015.

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Mission Relevance
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• The NASA Science Plan highlights the following ideas
for future potential missions based on the science
goals highlighted in the Decadal Survey:

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Background– F region
• The F region of the ionosphere contains ionized gases at a
height of ~150–800 km above sea level
• The F region has the highest concentration of free electrons
and ions anywhere in the atmosphere.
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Image Credit: Encyclopedia Britannica,. The Day-And-Night Differences In The Layers Of Earth's Ionosphere.. 2012. Web. 13 Sept. 2015.

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Background– Spread-F Depletions
• Spread-F Depletions are the occurrence of the post sunset
and nighttime plasma irregularities in the F-region
ionosphere.
– Broad range scale sizes over several orders of magnitude from ~10 cm
to ~100 km.
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Image Credit: Chapagain, Narayan P. "Dynamics of equatorial spread F using ground-based optical and radar measurements." (2011).

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Science Traceability Matrix
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STM – Primary Science Objective
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Science Objectives
Measurement Requirements Instrument
Functional
Requirements
Mission Top
Level
RequirementObservables
Physical
Parameters
PRIMARY SCIENCE OBJECTIVE
An enhanced vertical drift
in the F-region after
sunset is characteristic of
the low-latitude
ionosphere, whose
intensity exhibits
seasonal, longitudinal,
solar cycle, ionospheric
storm and neutral forcing
dependencies.
Presence of indicators
(ion species) that the
constellation is straddling
a nightside enhancement Ion spectrum
measurement,
ion and electron
distributions,
electric and
magnetic fields
using four point
method
Wide dynamic range in energy
coverage from spacecraft potential to
40keV/e.
Maintaining
electrostatic &
electromagnetic
cleanliness of
measurement probes by
introducing constant
satellite spin.
Separate the major mass ion species,
that is those that contribute
significantly to total mass density to
confirm night time enhancement using
Ion Spectrometry
Presence of tangential
DC electric and normal
magnetic fields
Oscillating electric-field in three axis in
the range 50–8000 Hz and amplitude
range 10 mV m-1 to 1 V m-1.
Continuous active
spacecraft potential
control to maintain
ground voltage
Time delays between signals from four
different antenna elements(Electric
field measurement) on the same
spacecraft, with a time resolution of
110 s.
Measurements required
for a minimum of one
seasonal cycle
Presence of a vertical ion
flux over baseline value
Three-dimensional velocity distribution
of electrons in the energy range from
0.59 eV to 26.4 keV
Deployment of
measurement devices
on external booms to
reduce noise.

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STM – Secondary Science Objectives
• Objective 1: Spread-F depletions at lower latitudes can be
appropriately explained by the formation of a vortex in the
ambient plasma at sunset due to the different velocities of
plasma and neutral gasses.
• Objective 2: An enhanced vertical drift in the F-region dynamo
after sunset results in large scale electric fields that lead to
night time enhancements of global ionospheric storms.
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Image Credit: Chapagain,
Narayan P. "Dynamics of
equatorial spread F using
ground-based optical and radar
measurements." (2011).

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Measurement Technique
• Spatial correlation of a control volume
– Requires 4 non-coplanar points of measurement
• Simplest such configurations is a tetrahedron
• Similar non-coplanar measurement technique used
previously by ESA Cluster mission and NASA MMS
mission
13

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Mission Statement
The SAGAN mission will investigate the impact of
atmospheric wave energy, neutral forcing, and current
drifts on low & mid-latitude structuring in near-Earth
Plasma in the ionosphere. This will be achieved through
simultaneous in-situ measurements of electric and
magnetic flux, electron density, and ion composition in
the ionosphere using a microsatellite constellation.

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Mission Objectives
1) Insert four identical microsatellites into a
tetrahedron formation consisting of three distinct
orbital planes about a reference orbit
2) The four satellites must maintain a close tetrahedron
structure continuously throughout the orbit, for the
duration of the mission based on the reference orbit
3) Science payload data must be collected and stored
on-board for each microsatellite
4) The science data from each microsatellite shall be
returned back to the ground station and archived.

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Mission Success Criteria
Source Mission Success Criteria Minimum Full
MO - 1 Each satellite shall achieve sucessful launch vehicle seperation and detumbling X
MO - 1 Establish communication link and perform health checks for all subsystems and payload for each satellite X
MO - 1 Perform necessary maneuver to achieve desired orbit for each satellite X
MO - 2 Relative distances during the close approach tetrahedron must be less than 150 km X
MO - 3 Each satellite shall achieve an angular rotation rate of 3 RPM to achieve electromagnetic cleanliness for the payload X
MO - 3 Coherent measurements of vertical drift in the F-region at the day/night terminator must be measured for 6 months X
MO - 4 6 months of payload data from all payload instruments on each satellite must be transmitted to the ground station X
MO - 3 Coherent measurements of vertical drift in the F-region at the day/night terminator must be measured for 24 months X X
MO - 3 One coherent measurement of spread-F depletions post-sunset must be measured X X
MO - 3
One coherent measurement of vertical depletions and large scale current drifts in the F-region during an ionospheric
storm at the day-night terminator must be measured
X X
MO - 4 24 months of payload data from all instruments on each satellite must be transmitted to the ground station X X

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Concept of Operations

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Trajectory Design
• Orbit plane 1 (“reference orbit”): 550 km circular orbit inclined at
30°. Contains Sat #1 and Sat #2 phased 0.85° apart.
• Orbit plane 2: Eccentric to reference orbit (~535 km by 565 km
altitude) with same inclination and period. Contains Sat #3 initially
phased 0.42° from Sat #1
• Orbit plane 3: Eccentric to reference orbit (~535 km by 565 km
altitude) with 30.2° inclination and same period. Contains Sat #4
initially phased 0.60° from Sat #1
Orbital Parameters Satellite 1 Satellite 2 Satellite 3 Satellite 4
Altitude at Perigee (km) 550 550 535 535
Altitude at Apogee (km) 550 550 565 565
Semimajor Axis (km) 6928 6928 6928 6928
Eccentricity 0 0 0.0022 0.0022
Period (seconds) 5738.82 5738.82 5738.82 5738.82
Inclination (°) 30 30 30 30.2
Longitude of Ascending
Node at Launch (°)
280.31 280.31 280.31 280.31
Phase from Satellite #1 (°) 0 0.85 0.42 0.6

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Trajectory Design (cont.)
View gifs online

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Two Weeks in the Life

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• The spacecraft structure is an “extruded hexagon”
– Driven by solar array sizing and ESPA Grande size limits
– Consists of two main equipment platforms (MEP) with
unobstructed field of view and one internal main equipment
platform
• 4 bulkheads and 6 spars along vertices of hexagon.
– Machined from Aluminum 6061 T-651 and 0.25” thick
Structure

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Structure
• Optimal mass reducing design is an isogrid design
– Array of equilateral triangles to increase structural
performance
– 75% mass reduction for equivalent flat plate
– 24 #10-32 holes for bolting instruments to the structure

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Overall CAD
10 m
2.5 m
10 m
0.45 m
0.9m
0.99m

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Science Instruments
Science Payload
Fields
Energetic
Particles
Potential
Control

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Science Instrument – Fields Group
Electric Fields and Waves (EFW):
• Measure electric field and density fluctuations
• Four orthogonal booms carrying spherical sensors
deployed to 10 m in the spin plan
Spatio Temporal Analysis of Field Fluctuations (STAFF):
• Measures magnetic fluctuations up to 4 kHz
• Four 2.5m long boom-mounted three axis search coil
magnetometers and two data-analysis packages
Digital Wave Processing (DWP):
• Signal processing package responsible for coordinating
Fields operations and selecting operational modes
(burst and nominal)

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Science Instrument – Energetic Particles
Plasma Electron & Current Experiment (PEACE):
• Measures the distribution function of the electrons
in the energy range of 0.59 eV to 26.4 eV
Cluster Ion Spectrometer (CIS)
• Ionic plasma spectrometry package containing a Hot
Ion Analyzer (HIA) and time-of-flight ion Composition
and Distribution Function Analyzer (CODIF)
• CODIF measures the distributions of the major ions
• HIA designed for ion-beam and solar-wind
measurements

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Science Instrument – Potential Control
Active Spacecraft Potential Control (ASPOC)
• Active charge neutralization device equipped with ion emitters of the
liquid-metal ion-source type
• Mitigates effects of surface charging by active charge neutralization

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Telecommunications
• High science data acquisition rate
• Two ground station locations: Melbourne, Florida &
Brisbane, Australia
IRIS Transponder V2 (NASA JPL)
– X-Band (Rx/Tx)
– Scalable RF Output Power
– Radiation Tolerant
– BPSK Modulation with Convolution
R=1/4, K=7 & R.S. (255,223)
ViaSat X-Band Ground Station
– 5.4m reflector
– 31.5 dB Gain
– Automated X-Y tracking
– Multiple Spacecraft Per
Antenna (MSPA) capability
AntDevCo Medium Gain
X-Band Patch Antenna
– 16.5 dB Gain
– 30° full
beamwidth

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Attitude Determination and Control
• Attitude knowledge (0.1°) requirement
• High torque (0.11 Nm) & angular momentum (24 Nms)
requirement
• Required Slew Rate & Angular Acceleration for a single ground
station pass

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Attitude Determination and Control

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Propulsion
• Driven from ΔV budget and science instrument design
– Can’t have artificial electric fields present near the
measurement probe
• Aerojet Rocketdyne MPS-230
– Modified to use AF-M315E “Green” Monopropellant
– 22 N primary thruster with four 1 N ADCS thrusters

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Guidance, Navigation, and Control
• GNSS Satellites used for precise inertial position and
velocity for each satellite
– Auto/cross-correlation measurement technique
– Collision Risk
NovAtel GPS-703-GGG
• High vibration variant available
• GPS, GLONASS, Galileo, BeiDou
signal reception
NovAtel OEM 615
• L1/L2 precise point positioning
(PPP) < 1.5 m
• Handles ionospheric effect
through linear combination of
L1 and L2 carrier phase

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Electrical Power System
• Science, Comms, ADCS components drive large power demands
• Solar Incidence Angle varies for spinning spacecraft
– Average projected area used (0.56 m2)
• Solar Aspect Angle within ±15°
Clyde Space FLEX EPSClyde Space Batteries
• 6 X 30 Whr
batteries, sync with
PDM board
• Built in heaters
MMA Body Mounted Solar Panels
• 1 custom sized panel per face
on hexagonal structure
• Triple Junction, 28.3% efficiency
• Dimensions: 0.43 m X 0.7 m
• 92 W Peak Sunlight Power

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C&DH and FSW
• Proton 200k DSP Processor Board
– 2 flight computers in cold-string configuration
• NASA core Flight System (cFS) Application Suite

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Thermal
Low (°C) High (°C)
ASPOC -25 65
EFW -60 100
STAFF -60 100
DWP -55 150
CIS -20 40
PEACE -40 85
IRIS Transponder V2 -25 75
AntDevCo X-Band MGA -40 70
RSI 45 Reaction Wheels -20 65
BCT FleXcore with IMU -40 85
BCT Nano Star Tracker -40 70
Adcole Mini Spinning Sun Sensor -40 70
Rocketdyne MPS-230 -22 50
Propellant Tank -40 80
NovAtel OEM615 Reciever -40 85
NovAtel GPS-703-GGG Antenna -40 85
C&DH Proton 200k Lite Processor Board -20 40
Clyde Space FLEX EPS -40 85
CS 30 Whr Battery -10 50
MMA Body Mounted Solar Panels -100 150
Overall Thermal Range -10 40
EPS
Telecom
Operating Thermal Range
Component NameSubsystem
Science Payload
Propulsion
ADCS
GNC
Aluminum paint coating
Absorptivity = 0.4
Emissivity = 0.53

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CAD – Exploded View

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Master Equipment List
Subsystem Component Name Quantity
CBE Mass
(kg)
TRL Contingency
MEV
Mass (kg)
Subsystem
MEV Mass
(kg)
ASPOC 1 1.85 6 25% 2.31
EFW 1 3.16 6 25% 3.95
STAFF 1 3.03 6 25% 3.79
DWP 1 2.05 6 25% 2.56
CIS 1 10.79 6 25% 13.49
PEACE 1 5.49 6 25% 6.86
Custom Hexagonal Structure 1 48.24 7 20% 57.89
Side Body Radiation Shielding 6 0.81 7 20% 5.85
Top/Bottom Radiation Shielding 2 1.33 7 20% 3.19
IRIS Transponder V2 1 1.10 7 20% 1.32
AntDevCo X-Band MGA 1 0.30 8 10% 0.33
RSI 45 Reaction Wheels 2 7.70 7 20% 18.48
BCT Nano Star Tracker 2 0.35 8 10% 0.77
BCT FleXcore with IMUs 1 0.85 7 20% 1.02
Adcole Mini Spinning Sun Sensor 1 0.25 7 20% 0.30
NovAtel OEM615 Reciever 1 0.02 8 10% 0.03
NovAtel GPS-703-GGG Antenna 1 0.50 8 10% 0.55
Rocketdyne MPS-230 1 0.40 6 25% 0.50
Propellant Tank 1 1.00 6 25% 1.25
Thermal Bright Aluminum Paint 1 4.88 8 15% 5.61 5.61
C&DH Proton 200k DSP Processor Board 2 0.20 7 20% 0.48 0.48
Clyde Space 3G FLEX EPS 1 0.17 7 20% 0.21
CS 30 Whr Battery 6 0.26 8 10% 1.72
MMA Body Mounted Solar Panels 6 1.50 6 25% 11.25
143.7
8.39
300
152.1
97%
32.96
1.65
1.75
13.17
0.58
66.93
20.57
Science
Payload
Telecom
Propulsion
EPS
GNC
Structures
ADCS
MEV Dry Mass (kg)
Propellent Mass (kg)
MPV Mass (kg)
MEV Wet Mass (kg)
Margin (%)
MEV Wet Mass
152.1 kg
Margin
97%

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Subsystem Component Quantity Contingency
CBE Power
Draw (W)
Duty
Cycle
CBE Power
Draw (W)
Duty
Cycle
CBE Power
Draw (W)
Duty
Cycle
CBE Power
Draw (A)
Duty
Cycle
ASPOC 1 15% 2.7 0 2.7 0 2.7 0 2.7 1
EFW 1 15% 3.7 0 3.7 0 3.7 0 3.7 1
STAFF 1 15% 2.96 0 2.96 0 2.96 0 2.96 1
DWP 1 15% 1.41 0 1.41 0 1.41 0 1.41 1
CIS 1 15% 10.64 0 10.64 0 10.64 0 10.64 1
PEACE 1 15% 8.457 0 8.457 0 8.457 0 8.457 1
IRIS Transponder V2 1 15% 17 0.2 17 1 17 1 17 0.01
AntDevCo X-Band MGA 1 15% 10 0.2 10 1 10 1 10 0.01
RSI 45 Reaction Wheels 2 15% 7 0.5 7 0 9 1 7 0
BCT Nano Star Tracker 2 10% 1.2 1 1.2 1 1.2 1 1.2 1
BCT FleXcore with IMU 1 10% 1.05 1 1.05 1 1.05 1 1.05 1
Adcole Mini Spinning Sun Sensor 1 10% 0.5 1 0.5 1 0.5 1 0.5 1
NovAtel OEM615 Reciever 1 10% 1.2 0.5 1.2 1 1.2 1 1.2 1
NovAtel GPS-703-GGG Antenna 1 10% 0.648 0.5 0.648 1 0.648 1 0.648 1
Propulsion Rocketdyne MPS-230 1 20% 28 0 28 1 28 0 28 0
C&DH Proton 200k DSP Processor Board 1 10% 2.3 1 2.3 1 2.3 1 2.3 1
CS 3G FLEX EPS 1 10% 0.5 1 0.5 1 0.5 1 0.5 1
CS 30 Whr Battery 6 10% 0.05 1 0.05 1 0.05 1 0.05 1
10%10%
128% -30% 18%
57.7 57.7 57.7
25.3 81.9 48.9
57.7
67.7
-15%
10% 10%
Science Mode
Science
Payload
Communication
Mode
61.523.0 74.4 44.4
Telecom
EPS
Safe Mode Maneuver Mode
GNC
ADCS
Margin
System Level Contingency
MEV Power Consumption (W)
Orbit Average Power Production (W)
Total MEV Power Consumption
Power Budget

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Battery Analysis – Science Mode
Battery Specs (Clyde Space 30Whr):
• Quantity: 6
• Capacity: 22.5 Ahr
• Rated DoD: 30% = 5.25 Ahr
(tested for 5000 cycles, 35000 expected)
Results:
• 13% Battery Margin above DoD (0.65 Ahr)
• 33% Power Production Margin
• 11,000 cycles expected for 2 years

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Battery Analysis – Comms. Mode
Battery Specs (Clyde Space 30Whr):
• Quantity: 6
• Capacity: 22.5 Ahr
• Rated DoD: 30% = 5.25 Ahr
Results:
• 16% Battery Margin above DoD (1.05 Ahr)
• Maneuver mode only during sunlight

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Thermal Energy Balance
• 1 Node Model
• Thermal Range between -10°C to 40°C

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Link Budget
Item Symbol Units Downlink Uplink
Frequency f GHz 8.45 8.45
Frequency f Hz 8.45E+09 8.45E+09
Transmitter Power (RF Output
Transmit Power)
P Watts 4 20
Transmitter Power (RF Output
Transmit Power)
P dBW 6 13
Transmitter Line Loss Ll dB -2 -2
Transmitter Antenna Gain Gt dBi 13.5 31.0
Equivalent Isotropic Radiated Power EIRP dBW 17.5 42.0
Propagation Path Length S m 1709926 1518364
Speed of Light c m/s 299792458 299792458
Free Space Path Loss Ls dB -175.6 -174.6
Propagation and Polarization Loss La dB -2 -2
Receive Antenna Pointing Loss Lpr dB -2 -2
Receive Antenna Gain Gr dBi 31.0 13.5
System Noise Temperature Ts K 300.0 80.0
System Noise Temperature Ts dBK 24.8 19.0
R bps 8346862 2000000
R kbps 8347 2000
R Mbps 8.35 2
Symbols per Byte - - 2 2
Eb/No Eb/No dB 6.5 26.5
Carrier-to-Noise Density Ratio C/No dB-Hz 72.7 86.5
Required Eb/No Req Eb/No dB 1.5 1.5
Implementation Loss - dB -1 -1
Margin - dB 4 24
Data Rate
Downlink
8.35 Mbps
4 dB Margin
Uplink
2 Mbps
24 dB Margin

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Data Return Strategy
• Strategy depended on high science data volume
• Average Downlink analysis performed via STK for 2 years for one
satellite
• Overall Mission Operations Data Volume Analysis
Item Nominal Mode Burst Mode
Data Acquisition Time (min/orbit) 63.6 32.0
Data Volume (Mb/orbit) 46.64 190.2
Total Data Volume (Mb/orbit)
Science Data Volume Analysis
236.85
Avg # Passes per Day 14
Avg Pass Duration (minutes) 11.4
Average Gap Between Overpass (minutes) 91.6
Average Blackout period per day (hrs) 10
Downlink Data Volume Capability (Gb/orbit) 5.3
Downlink Data Volume Analysis
Number of Orbits Per Day 15
Mission Science Mode Percentage 93.5%
Science Mode Time Per 2 Weeks (days) 13
Mission Communication Mode Percentage 6.5%
Communication Mode Time Per 2 Weeks 1
Transmission Data Volume (Gb/2 weeks) 73
Science Data Volume (Gb/2 weeks) 46.7
Margin 56%
Mission Operations Data Volume Analysis

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Delta-V Budget
Maneuver Type Description ΔV (m/s) Margin Total ΔV (m/s)
Plane Change (0.2°) 26.477 25% 33.097
Max Phasing Burn (0.85°) 11.636 25% 14.545
Radial Impulse Burn 15.886 25% 19.858
Altitude Maintainence 550 km - 2 years 22.532 25% 28.165
Attitude Control
Desaturation of Reaction
Wheels - 2 years
10.649 25% 13.311
De-Orbit Drag Deorbit 0.000 0% 0.000
108.975
10%
119.872Total ΔV (m/s):
Maneuvers to Achieve
Initial Orbit
Sum of ΔV (m/s):
Overall Margin:
• Plane Change: ΔV perpendicular to orbit plane
• Phasing Burn: Elliptical transfer into same orbit
• Radial Impulse Burn: Flight path angle adjustment
• Altitude Maintenance: Ballistic coefficient and orbital
parameters give ΔV needed per orbit

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Management
Summary
Management Outline

49
Motivation/
Background
Mission
Overview
Flight System
Systems
Engineering
Program
Management
Summary
Program Schedule
Key Milestones
Phase A: Mission &
Systems Definition
Phase B:
Preliminary Design
Phase C: Final
Design &
Fabrication
Phase D: System
Integration, Testing,
and Verification
Phase E: Operations
Phase F: Closeout
2023
Q1 Q2 Q3 Q4
2022
Q1 Q2 Q3 Q4
2021
Q1 Q2 Q3 Q4
Project
Phase
2015 2016
Q4
2017 2018 2019
Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4 Q1 Q2 Q3
2020
Q1 Q2 Q3 Q4
SRR MDR PDR ARR MRRORRTRR PLAR CERR DR
SDR – System Requirements Review
MDR – Mission Design Review
PDR – Preliminary Design Review
ARR – Assembly Readiness Review
TRR – Test Readiness Review
ORR – Operational Readiness Review
MRR – Mission Readiness Review
PLAR – Post Launch Assessment Review
CERR – Critical Event Readiness Review
DR – Decommissioning Review

50
Motivation/
Background
Mission
Overview
Flight System
Systems
Engineering
Program
Management
Summary
Risk Analysis
Consequence Qualitative Definition
1 Minimal or no impact to mission
2 Small reduction in mission return
3 Cannot meet full mission success
4 Cannot meet minimum mission requirements
5 Mission catastrophe
Likelihood Qualitative Definition Probablility Range
1 Very Low < 1%
2 Low 1-5%
3 Moderate 5-15%
4 High 15-30%
5 Very High > 30%

51
Motivation/
Background
Mission
Overview
Flight System
Systems
Engineering
Program
Management
Summary
Risk Analysis (cont.)
1 2 3 4 5
5
4
D
3 D' C' C
2 B A
1 B' A'
Likelihood
Consequence
Risk
Unmitigated
Likelihood
Unmitigated
Impact
Handling Method
A
Collision of two or more
satellites while crossing orbits
2 5
Higher fidelity GPS and development of
collision detection algorithm
B
Failure to complete design by
launch date
2 4
Increased time margins to schedule and
earlier testing of lower TRL components
C Failure of ADCS actuators 3 4 Additional propellant for ADCS thruster
Reliability and environment qualifications
testing early on in Phase C
Redundancy of internal mechanisms
4 4D
Failure of boom deployment
mechanism

52
Motivation/
Background
Mission
Overview
Flight System
Systems
Engineering
Program
Management
Summary
Cost Estimation Approach
• Parametric Cost Estimation
– Small Satellite Spacecraft Model (SSCM)
• Accounts for both recurring and non-recurring costs
• Bottom-Up Method
– Exact cost values for hardware used
– More specific to our mission
• 4 flight unit + 1 engineering unit
– Can better account for development costs for science
instruments

53
Motivation/
Background
Mission
Overview
Flight System
Systems
Engineering
Program
Management
Summary
Cost Analysis – Bottom Up
WBS Element Quantity Years Cost ($k)
Component
Margin (%)
Final Cost ($k)
Total Hardware Cost 1 - 1,052.10$ Included 1,052.10$
IA&T N/A - 100.00$ 25% 125.00$
Total Hardware Cost 4 - 7,963.06$ Included 7,963.06$
IA&T N/A - 600.00$ 25% 600.00$
ASPOC 4 - 2,845.80$ Included 2,845.80$
EFW
STAFF
DWP 4 - 500.00$ Included 500.00$
CIS 4 - 865.00$ Included 865.00$
PEACE 4 - 875.00$ Included 875.00$
Development and IA&T 4 - 8,000.00$ 25% 10,000.00$
Principal Investigator 1 8.5 4,250.00$ 10% 4,675.00$
Mission Design Engineer 16 5 16,000.00$ 10% 17,600.00$
Mission Ops. Engineer 8 2 2,880.00$ 10% 3,168.00$
Ground Support Engineer 3 2.5 1,500.00$ 10% 1,650.00$
Science Personnel 5 8.5 6,375.00$ 10% 7,012.50$
Management 2 8 2,400.00$ 10% 2,640.00$
Launch Opportunity 1 - 6,250.00$ 10% 6,875.00$
Ground Support Equipment 2 - 2,000.00$ 25% 2,500.00$
72,997.35$
25%
91,246.69$
91,246,692.23$
-
Program Level
Flight Support
Cost ($K)
System Margin
Ground Equipment
Total Cost ($K)
Total Cost ($)
2,025.00$ Included 2,025.00$4
Spacecraft - Engineering Unit
Science Payload
Spacecraft - Flight Unit
Total Cost
$91,246,692.23

54
Motivation/
Background
Mission
Overview
Flight System
Systems
Engineering
Program
Management
Summary
Descope Options
• Option #1: Replace EFW and STAFF
- Lower spatial and temporal resolution
- Hinder observing extrema of range of physical phenomenon
+ Development cost savings of roughly $2.5 million
+ Improves margin for science instrument development
• Option #2: Move from X-band to S-band transmission
- Lower downlink rate implies longer downlink time
- Decreased time spent in science mode
+ Cost savings in purchase and maintenance of ground station
+ Relaxes ADCS slew requirements due to greater beam width

Motivation/
Background
Mission
Overview
Flight System
Systems
Engineering
Program
Management
Summary
55
Conclusion
• SAGAN mission will help answer fundamental science
questions deemed relevant by the RFP
• Highest spatial/temporal resolution of near Earth plasma
ever attempted
• Suite of instruments and subsystem design fully closes the
mission design outlined by the requirements
• Within cost cap of typical NASA Small Satellite Mission
– ~$92 million cost (36% less)
55

Motivation/
Background
Mission
Overview
Flight System
Systems
Engineering
Program
Management
Summary
56
Acknowledgements
• Dr. Glenn Lightsey – Professor of Aerospace Engineering
• Terry Stevenson – Graduate Teaching Assistant
• Jason Swenson – Graduate Teaching Assistant
• Dr. Carol Paty – Professor of Earth and Atmospheric
Sciences
• Dr. Sara Spangelo – Systems Engineer, JPL
• Dr. Morris Cohen – Professor of Electrical and Computer
Engineering
56

“Somewhere, something incredible is waiting to be known.”
-Carl Sagan

Motivation/
Background
Mission
Overview
Flight System
Systems
Engineering
Program
Management
Summary
Background – Atmospheric Waves
• Atmospheric waves are global-scale waves
– Excited by regular solar differential heating (solar thermal tides)
– The gravitational tidal force of the moon (gravitational tides)
– Neutral forcing.
58
Altitude-Latitude structure of electron density (in MHz) due to TW3 type tidal waves

Motivation/
Background
Mission
Overview
Flight System
Systems
Engineering
Program
Management
Summary
Background – Ionospheric Storms
• Ionospheric Storms: Disturbances due to solar activity, when
affecting the ionosphere, are ionospheric storms
– Tend to generate large disturbances in ionospheric density distribution,
total electron content, and the ionospheric current system
59
Image Credit: Chapagain, Narayan P. "Dynamics of equatorial spread F using ground-based optical and radar measurements." (2011).
Measurement of the ionosphere's total electron content (TEC) using GPS Mapping

MMS Mission
• Studying magnetic field turbulence measurements at the
bow shock
• Studying magnetic reconnection
• Highly elliptical orbit at ~10 earth radii away
• 4 satellites are a measurement technique for spatial and
temporal autocorrelation
12/4/2015 60

Level 1 Requirements
MDR Mission Design Requirements Source Verification
MDR - 1 The mission duration shall last at least 6 months with a goal of 24 months in the initial orbit MO - 3 Inspection
MDR - 2
Satellites #1 and #2 must be placed in the reference orbit (circular, 550 km altitude, 30° inclination,),
phased 0.85° degrees apart
MO - 1 Analysis
MDR - 3
Satellite #3 must be eccentric to the reference orbit such that it maintains the same period as Satellite #1
with altitude at perigee of 535 km, altitude at apogee of 565 km, and phased 0.42° from Satellite #1
MO - 1 Analysis
MDR - 4
Satellite #4 must be positively inclined to the reference constellation orbit by 0.2 degrees and eccentric to
the reference orbit such that it maintains the same period as Satellite #1 with altitude at perigee of 535 km,
altitude at apogee of 565 km, and phased 0.60° from Satellite #1
MO - 1 Analysis
MDR - 5
Once in tetrahedron formation, each satellite shall maintain the orbital parameters of period, eccentricity,
inclination, and phase from Satellite #1 for the duration of the mission
MO - 3 Analysis
MDR - 6 End of life drag deorbit plan must be initiated in order to meet the 25 year deorbit plan set forth by NASA MO - 2 Analysis
FSR Flight System Requirements Source Verification
FSR - 1 Each satellite must be able to withstand the launch vehicle environment MO - 1 Analysis
FSR - 2 Each satellite must be able to survive during operations in space for mission duration MO - 3 Testing
FSR - 3
Each satellite must maintain spacecraft attitude relative to other satellites and maintain correct orbital
parameters for the tetrahedron formation
MO - 2 Analysis
FSR - 4 Each satellite must have the ability to store and relay payload and state data back to ground station MO - 3 Testing

Level 1 Requirements (cont.)
SPR Science Payload Requirements Source Verification
SPR - 1
The science payload must cover a wide range of energies, from spacecraft potential to 40 keV/e for ion
spectroscopy measurements
MO - 3 Testing
SPR - 2
The science payload shall measure a wide dynamic range in energy coverage from spacecraft potential to
40 keV/e.
MO - 3 Testing
SPR - 3
The science payload must measure electric-field in the frequency range of 50–8000 Hz and amplitude range
10 mV m-1
to 1 V m-1
.
MO - 3 Testing
SPR - 4 The science payload shall measure with a time resolution of 1.10E-5 s. MO - 3 Testing
SPR - 5
The science payload must measure three-dimensional velocity distribution of electrons in the energy range
from 0.59 eV to 26.4 keV
MO - 3 Testing
SPR - 6 The science payload must measure three-axis magnetic fluctuations up to 4 kHz MO - 3 Testing
SPR - 7 The science payload must maintain electromagnetic cleanliness for the duration of the mission MO - 3 Testing
SPR - 8 The science payload must actively maintain ground potential for optimum measurments MO - 3 Testing
GSR Ground System Requirements Source Verification
GSR - 1 The ground system shall downlink science data and telemetry from all satellites with a margin of atleast 2 dB MO - 3 Testing
GSR - 2 The ground system must uplink commands to all satellites with a margin of atleast 15 dB MO - 3 Testing
GSR - 3 The ground system shall comprehensively archive all data received MO - 3 Testing
GSR - 4
The mission personnel must perform all necessary mission operations for the lifetime of the mission and
transfer all science data to science personnel
MO - 3 Testing

Level 2 Requirements
ADCS ADCS Requirements Source Verification
ADCS - 1
ADCS maintain spin stabilization in inertial space during all modes of
operation
FSR - 2 Analysis
ADCS - 2
ADCS shall be able to reorient the spacecraft within a full range of
motion for orbit insertion and maintenance
FSR - 3 Analysis
ADCS - 3
ADCS shall maintain a solar aspect angle of 90 ± 15° during science
(nominal) mode for power acquisition
FSR - 2 Analysis
ADCS - 4
ADCS shall maintain a slew rate of at least 0.74°/s with
a ±15° transverse pointing accuracy for periods of at least 773 s in
order to relay data to the ground stations
FSR - 4 Analysis
ADCS - 5
ADCS shall maintain a pointing knowledge of 0.1° for attitude
determination during science mode
FSR - 3 Testing
ADCS - 6
ADCS shall provide each satellite with a rotation rate of 3 rpm in
order to maintain electromagnetic cleanliness
SPR - 7 Analysis
ADCS - 7
ADCS thrusters shall be capable of dumping additional momentum
(relative to the nominal spin rate) over the period of the mission
FSR - 4 Analysis
ADCS - 8
ADCS shall reorient spacecraft during the end of life operations to
deorbit within 25 years
MDR - 6 Analysis
TCS Thermal Control System Requirements Source Verification
TCS - 2
TCS must maintain a temperature range between -10° C and 40° C
at all times bfore and after deployment fromlaunch vehicle
FSR - 2 Analysis

CDH Command & Data Handling Requirements Source Verification
CDH - 1
C&DH shall be able to process spacecraft telemetry at a minimum
rate of 5 Hz
FSR - 4 Testing
CDH - 2 C&DH shall handle subsystem control at a minimum rate of 10 Hz FSR - 2 Testing
CDH - 3
C&DH shall store at least 240 Mb of data per orbit and at least 215
orbits worth of data
FSR - 4 Testing
CDH - 5 C&DH shall provide data interfaces for each subsystem FSR - 4 Testing
EPS Electrical Power System Requirements Source Verification
EPS - 1
EPS shall provide 25.3 Watts to the spacecraft bus and 0 Watts to
the payload during safe mode
FSR - 2 Testing
EPS - 2
EPS shall provide 14.63 Watts to the spacecraft bus and 34.6
Watts to the payload during science mode
FSR - 2 Testing
EPS - 3
the payload during communications mode
FSR - 2 Testing
EPS - 4
the payload during thrust mode
FSR - 2 Testing
EPS - 6
EPS shall store 22.5 Amp-Hrs of electrical power during mission
lifetime
FSR - 2 Testing/Analysis

PROP Propulsion System Requirements Source Verification
PROP - 1
Propulsion must provide 34 m/s of ΔV in order to achieve a 0.2°
plane change
MDR - 2 Testing
PROP - 2
Propulsion must provide 20 m/s of ΔV in order to create the desired
eccentric orbit of 535 km by 565 km altitude
MDR - 2 Testing
PROP - 3
Propulsion must provide 29 m/s of total ΔV over 24 months to
perform station keeping
MDR - 2 Testing
PROP - 4
Propulsion must provide 15 m/s of total ΔV per satellite in order to
perform phasing maneuvers
MDR - 2 Testing
PROP - 5
Propulsion must provide 14 m/s of total ΔV per satellite in order to
desaturate the reaction wheels
MDR - 2 Testing
COMMS Communication Systems Requirements Source Verification
COMMS - 1
Comms shall uplink at a minimum rate of 2 Mbps per orbit with a 24
dB margin to ground station
GSR - 2 Testing
COMMS - 2
Comms shall downlink all data captured in 197 within 14 ground
station passes with a 4 dB margin
GSR - 1 Testing
COMMS - 3
Comms shall allow uplink and downlink occur only to the assigned
ground stations
FSR - 4 Analysis

STRUCT Structures Requirements Source Verification
STRUCT - 1 The volume of each satellite must be under 1 m
3
FSR - 1 Inspection
STRUCT - 2 The mass of each satellite must be under 300 kg FSR - 1 Inspection
STRUCT - 3
Structure must survive a dynamic load equivalent to 4.55 g's during
launch
FSR - 1 Analysis
STRUCT - 4
Structure must be stiff enough to survive a 20-45Hz oscillation along
all axis with a safety factor of 11 during launch
FSR - 1 Analysis
STRUCT-5 Structure must facilitate body mounted solar arrays FSR-2 Analysis
STRUCT - 6
Structure must be able to maintain internal temperature and radiation
levels up to 1 kRad at all time.
FSR - 2 Analysis
STRUCT - 7
The shape of the structure must be optimized to maximize projected
area for solar energy and internal volume within the volume
FSR - 2 Analysis
GNC Guidance, Navigation, and Control Requirements Source Verification
GNC - 1
GNC shall acquire intertial position vectors within a 1.2 m precision
for formation flying
FSR - 3 Analysis
GNC - 2
GNC shall acquire intertial velocity vectors within a 10 m/s precision
for formation flying
FSR - 3 Analysis
GNC - 3
GNC shall maintain the commanded orbit track of each satellite with
an absolute position error of no more than 1 km
MDR - 3 Analysis

ADCS – Slew Phasing
12/4/2015 67

Cost Estimation (MEL)
Subsystem Component Name Quantity TRL Cost FY'15 ($)
Contingency
(%)
Component Cost
($)
Subsystem Cost
($)
ASPOC 1 6 569,160.00$ 25% 711,450.00$
EFW 1 6
STAFF 1 6
DWP 1 6 100,000.00$ 25% 125,000.00$
CIS 1 6 173,000.00$ 25% 216,250.00$
PEACE 1 6 175,000.00$ 25% 218,750.00$
Custom Hexagonal Structure 1 7 7,650.00$ 20% 9,180.00$
Side Body Radiation Shielding 6 7 3,750.00$ 20% 4,500.00$
Top/Bottom Radiation Shielding 2 7 1,250.00$ 20% 1,500.00$
IRIS Transponder V2 1 7 250,000.00$ 20% 300,000.00$
AntDevCo X-Band MGA 1 8 36,000.00$ 10% 39,600.00$
RSI 45 Reaction Wheels 2 7 168,000.00$ 20% 201,600.00$
BCT Nano Star Tracker 2 8 136,000.00$ 10% 149,600.00$
BCT FleXcore 1 7 110,000.00$ 20% 132,000.00$
Adcole Mini Spinning Sun Sensor 1 7 15,000.00$ 20% 18,000.00$
NovAtel OEM617 Reciever 1 8 2,775.62$ 10% 3,053.18$
NovAtel GPS-703-GGG Antenna 1 8 1,595.00$ 10% 1,754.50$
Rocketdyne MPS-230 1 6
Propellant Tank 1 6
Propellant (AF-M315E) 1 6 3,500.00$ 25% 4,375.00$
C&DH Proton 200k Lite Processor Board 1 7 75,000.00$ 20% 90,000.00$ 90,000.00$
Clyde Space 3G FLEX EPS 1 7 13,500.00$ 20% 16,200.00$
CS 30 Whr Battery 3 8 5,400.00$ 10% 5,940.00$
MMA Body Mounted Solar 6 6 436,677.96$ 25% 545,847.45$
Thermal Bright Aluminum Paint 4 8 94,148.00$ 10% 103,562.80$ 103,562.80$
Total Cost: 3,846,137.93$
Structures
ADCS
15,180.00$
501,200.00$
EPS
Propulsion
353,380.00$
Telecom
GNC
Science
Payload
405,000.00$
567,987.45$
25%
25%
506,250.00$
441,725.00$
1,777,700.00$
339,600.00$
4,807.68$
446,100.00$

Cost Estimation (Parametric - SSCM)
WBS Element Driver - Subsystem Mass (kg) Cost ($k)
Structures 66.93 5,837.06$
Telecom 1.65 595.12$
ADCS & GNC 21.15 6,994.21$
Propulsion 1.75 95.08$
C&DH 0.48 685.84$
EPS 13.17 4,709.98$
Thermal 5.61 514.39$
WBS Element Driver - S/C Bus Cost ($K) Cost ($k)
Science Instruments 77,726.69$ 31,090.68$
Integration, Assembly & Test 19,431.67$ 2,701.00$
Program Level 19,431.67$ 4,449.85$
Launch & Orbital Operations 19,431.67$ 1,185.33$
Ground Support Equipment 19,431.67$ 1,282.49$
Total Cost - 1 Satellite ($K) 36,823.02$
Total Cost - 1 Satellite ($) 36,823,021.58$
Total Cost - 4 Satellites($K) 118,436.05$
Total Cost - 4 Satellites ($) 118,436,050.95$
Spacecraft
Payload
Ground Equipment
Flight Support
Program Level
Spacecraft Integration, Assembly, and Test

Propulsion Trade Study
12/4/2015 72
Specifications Weight AR MR-111C AR MPS-230 Airbus S10 AR R-6D
Propellant Type - Mono Mono Bi Bi
Dry Mass of Thruster 1 S - - -
Propellant Toxicity 2 S + S S
Isp 3 S + + +
Nominal Thrust 3 S + + +
Propellant Mass Required for ΔV 3 S + + +
Power Requirement 2 S - - S
Attitude Control Thrusters 3 S + - -
TRL Level 2 S - S S
Total + 14 9 9
Total - 5 6 4
Total S 0 4 6

Data Return Strategy – STK Ground Pass
12/4/2015 74

Data Return Strategy – STK Ground Pass
Access # Access Start (UTCG) Access End (UTCG) Duration (sec) Asset Full Name Data Volume (Gb)
428 2/1/21 7:10 2/1/21 7:21 666 Brisbane 5.56
429 2/1/21 8:50 2/1/21 9:03 772 Brisbane 6.44
430 2/1/21 9:32 2/1/21 9:38 339 Melbourne 2.83
431 2/1/21 10:31 2/1/21 10:44 780 Brisbane 6.51
432 2/1/21 11:10 2/1/21 11:22 704 Melbourne 5.88
433 2/1/21 12:13 2/1/21 12:26 779 Brisbane 6.50
434 2/1/21 12:51 2/1/21 13:04 777 Melbourne 6.49
435 2/1/21 13:55 2/1/21 14:07 775 Brisbane 6.47
436 2/1/21 14:32 2/1/21 14:45 780 Melbourne 6.51
437 2/1/21 15:36 2/1/21 15:48 684 Brisbane 5.71
438 2/1/21 16:14 2/1/21 16:27 781 Melbourne 6.52
439 2/1/21 17:21 2/1/21 17:25 185 Brisbane 1.55
440 2/1/21 17:55 2/1/21 18:08 761 Melbourne 6.35
441 2/1/21 19:37 2/1/21 19:48 612 Melbourne 5.11
78.4Total Data Volume (Gb):
12/4/2015 75

Solar Panel Sizing
Pe (W) 46
Pd (W) 91.9
Max Eclipse % 31.3%
Max Sunlight % 68.7%
Te (s) 2170.75
Td (s) 4757.35
Xe 0.6
Xd 0.8
Psa (W) 160.35
Mean Solar Flux (W/m2
) 1370
Solar Cell Efficiency (%) 28.30%
Po (W/m
2
) 387.71
Id 0.77
θ 15
PBOL (W/m2
) 288.36
PEOL (W/m2
) 285.49
Area - Sollar Array (m
2
) 0.5617
Solar Area Sizing
Base Edge (m) 0.43
Height (m) 0.7
Best Case Projected Area (m2
) 0.602
Worst Case Projected Area (m2
) 0.5213
Avg Projected Area (m2
) 0.5617
Minimum Solar Panel Face Sizing
12/4/2015 76

SAGAN Mission to Study Atmospheric Gradients and Neutral forcing in Near-Earth Plasma

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