FLARE Final Presentation (no animations)

FLARE
(Flyby Anomaly Research Endeavor)
• Project Manager:
– Amritpreet Kang
• Systems Engineer:
– Graeme Ramsey
• Chief Engineer:
– Jeffrey Alfaro
• Associate Engineers:
– Kyle Chaffin
– Anthony Huet
𝐾 =
2ω 𝑒 𝑅 𝑒
𝑐
= 3.099 x 10−6
*point mass orbital mechanics,
2D flyby visual

Executive Summary
• Team FLARE of the University of Texas at Austin has been tasked with confirming the flyby anomaly notably
experienced first by Galileo in 1990 followed by NEAR, Cassini, Messenger and Rosetta.
• The anomaly takes the form of an unaccounted for change in energy/velocity which takes place around periapse
of a hyperbolic planetary flyby during which their is a change in declination. The velocity anomalies vary by as
much as 13.5 mm/s from precisely modeled values.
• A phenomenological formula which relates the velocity discrepancy to a change in declination, excess velocity and
a constant scaling factor serves to guide a flyby trajectory corollary to the anomaly.
• Many causes have been conjectured, accounted for, or otherwise proved innocent (from atmospheric drag to
modifications to inertia). A thorough investigation of the navigation software and mathematical models used for
navigation by JPL uncovered two potential culprits (high order gravity terms and anisotropy of the speed of light).
• Team FLARE’s proposed design is an affordable CubeSat mission whose goal is to gather more data points on the
anomaly to corroborate its existence.
Speaker: Graeme Ramsey

Executive Summary
The primary benefit from this mission is filling in the gap of closest
approach left by most heritage missions and in the process prove whether the
anomaly truly exists. Furthermore, the data gained from FLARE would allow
further evaluation of the two most probable explanations of the anomaly.
This endeavor will lead to more accurate trajectory propagation
methods by further characterizing this anomalous perturbation. By those
standards, objects like Earth rendezvousing asteroids will be predictable to a
higher degree.

Anomaly Background
Speaker: Kyle Chaffin
Parameter GLL-I GLL-II NEAR Cassini Rosetta M'Ger Juno
Date 12/8/1990 12/8/1992 1/23/1998 8/18/1999 3/4/2005 8/2/2005
H (km) 960 303 539 1175 1956 2347
φ (deg) 25.2 -33.8 33 -23.5 20.2 46.95
λ (deg) 296.5 354.4 47.2 231.4 246.8 107.5
Vf (km/s) 13.74 14.08 12.739 19.026 10.517 10.389
V_inf (km/s) 8.949 8.877 6.851 16.01 3.863 4.056 9.91
DA (deg) 47.7 51.1 66.9 19.7 99.3 94.7
i (deg) 142.9 138.7 108 25.4 144.9 133.1
αi (deg) 266.76 219.35 261.17 334.31 346.12 292.61
δi (deg) -12.52 -34.26 -20.76 -12.92 -2.81 31.44 -14.2
αo (deg) 219.97 174.35 183.49 352.54 246.51 227.17
δo (deg) -34.15 -4.87 -71.96 -4.99 -34.29 -31.92 39.4
MSC (kg) 2497 2497 730 4612 2895 1086
ΔV_inf (mm/s) 3.92 -4.6 13.46 -2 1.8 0.02 0
σV_inf (mm/s) 0.3 1 0.01 1 0.03 0.01 2
Theoretical ΔV_inf (mm/s) 4.12 -4.67 13.28 -1.07 2.07 0.06 6.04
𝐾 =
2ω 𝑒 𝑅 𝑒
𝑐
= 3.099 x 10−6

Heritage Mission Data Acquisition Overview
Heritage missions navigation precision details [24-26, 26].
• Instruments used on heritage missions to obtain velocity data.
• With these instruments, NEAR measured the highest change in hyperbolic excess
velocity, whereas Juno measured no apparent change.
• Uniquely, Juno incorporated 50x50 and 100x100 gravitational modeling, leading to
mismatch between expected and apparent anomaly, in fact no apparent anomaly [36].
• Explanations of the flyby anomaly focus on modeling errors:
• Higher order gravity terms
• Anisotropy of the speed of light

Dominate Anomaly Sources: (JUNO) High Order Gravity
Terms and Anisotropy of the Speed of Light
• HOGT: Truncation in Earth’s geopotential model is actually a perturbation
capable of producing something detectable in real time comparable to the
predicted flyby anomaly [36].
• ASL: The flyby anomalies result from the assumption that the speed of
light is isotropic in all frames, but the speed of light is not invariant and
isotropic only with respect to a dynamical 3-space [44].
Speaker: Kyle Chaffin and Graeme Ramsey

Terms and Anisotropy of the Speed of Light
JUNO Doppler postfit residuals reconstruction (left) and deleted data (right) [36].

Simulated Doppler residuals from 7 mm/s anomaly with (left) and without (right) spin signature [36]

Position (top) and Velocity
(bottom) perturbations
incurred by modeling with
higher order gravity
models than 10X10 [36].
The order of this
perturbation is comparable
to that of the anomaly.

Phenomenological Formulae and Perturbation Magnitudes
𝐾 =
2ω 𝑒 𝑅 𝑒
𝑐
= 3.099 x 10−6
• Primary formula [1]
• Developed by: JPL (Anderson et al.
2008)
• Effective range: 500 to 2000 km
• Error: same as secondary formula
Magnitude of pertinent accelerations, courtesy
of a Portuguese mission proposal [39].
• Secondary formula [44]
• Developed by: Stephen Adler,
Institute for Advanced Study
• Similar range, see error table

Mission Drivers
Need Statement:
Evaluate whether the hyperbolic flyby anomaly is a consistent, repeatable phenomenon, or
an otherwise unaccounted for data artifact.
Goals:
Collect a quantity of at least 4 data points during hyperbolic flybys, showing repeatability of
the anomaly, and characterizing its effects.
Objectives:
Collect position, velocity, and acceleration data over the course of at least 4 hyperbolic
flybys from two spacecraft comparable to the data from the NEAR spacecraft Earth flyby.
Speaker: Jeffrey Alfaro

Primary Requirements
• [A] The system shall be capable of gathering the velocity profile during the inbound and
outbound legs of a hyperbolic flyby trajectory of Earth.
• [B] This project will provide at least 4 velocity profiles associated with the flyby
phenomenon in its projected lifetime.
• [C] The system shall be capable of tracking the velocity the satellite experiencing the
hyperbolic flyby anomaly during closest approach on the order of 0.1 mm/s accuracy.
• [D] The mission design shall perform velocity data collection on “paired” flybys (with
minimal separation) at the above mentioned accuracy (~0.1 mm/s), including coverage
throughout closest approach.

Secondary Requirements
• {A} The trajectory of the satellites during closest approach shall be monitored with GPS,
including back/side lobe GNSS tracking, the use of tens of ground stations and post
processing for added accuracy.
• {B} Confirmation of an anomalous DV shall be achieved via (Doppler effects) X-band
radio broadcasting during the flyby phases.
• {C} The error of Doppler velocity measurements shall be at maximum 0.5 mm/s.
• {D} The satellites will be constrained to a standard 3u/6u format.
• {E} The satellites will perform flybys with sufficient hyperbolic excess velocity and
change in declination to produce an anomaly of at least ±3 mm/s.
• {F} The altitude of periapse upon each flyby shall be between 500 and 2000 km, the
best fit range of the phenomenological formula.

Requirements Traceability
Items after [extra] are requirements that weren’t explicitly listed.
Traceability Matrix Relationship: X=direct O=indirect
Primary Mission
Primary V_inf accuracy 4 data pnts V accuracy Tandem sats Budget Mission Assurance Trajectory
Requirement [A] [B] [C] [D] [extra]
System GNSS {A} O X X
Doppler DSN {B} X O
Doppler error {C} X X X X
Sat Size {D} O X
Predicted anomaly {E} O X X
Altitude of periapse {F} X O X X

Constraints
• Projected satellite lifetime: 3 years.
– Radiation toll and propulsion capacity.
– 250-300 m/s DV corrections capable with 4u worth of hydrazine propulsion.
• (optional) cold gas attitude thrusters, desaturation maneuvers for reaction wheals.
– Medium to High TRL and rad hardened subsystem components only.
• Mission budget: $5mil before launch associated costs.
• Secondary payload considerations.
– Satellites must be compatible with a Planetary Systems CSD.
– Satellite mass: 12 kg CSD constraint. Max satellite volume: 6u.
• Launch window and parking orbit/exit trajectory characteristics.
– High eccentricity and inclination, Molniya type parking orbit (considering baseline trajectory.) [Primary ConOps]
– GTO parking orbit option. More ride sharing possibilities. [Secondary ConOps]
• Flyby characteristics must coincide with phenomenological formula.
• SHERPA must be compatible with the launch vehicle.

• CSD [4]
• 3.4 kg in mass
• X and Y dimensions: 26.34 cm and 15.75 cm
• Ejection plate force on payload from launch vibration: 0-191 N
• Ejection plate force on payload from spring ejection: 15.6-46.7 N
• Survival temperature extrema: -50 to 100 °C
• Operational temperature extrema: -45 to 90 °C
• Life: 50 door closures
• Payload [27]
• 12 kg max
• Tab lengths: X = 23.92 cm, Z = 36.5 cm
• Force from deployment switches, Z-axis: 5 N
• Friction from 4 sides contacting walls: 2 N
Capsulized Satellite Dispenser (CSD) Constraints

CSD payload specifications
courtesy of Planetary
Systems Corporation [27].

Capsulized Satellite Dispenser (CSD) Constraints
CSD specifications courtesy of Planetary Systems Corporation [4].

SHERPA Capabilities
SHERPA configuration (left) and capabilities (right), courtesy
of Spaceflight Inc. [3,25]

ConOps Intro: Post-Launch/Pre-Flyby Maneuver
SHERPA mounted on a primary payload of a Falcon 9 [25].

SHERPA deployment from Falcon 9 payload section [3].

SHERPA Rideshare potential [3].

SHERPA 6U CubeSat Deployment via a CDS [4].

ConOps A: Repeat tandem flybys of Earth
1. Launch as a secondary payload, highly
inclined.
2. SHERPA second stage provides
hyperbolic excess velocity for FLARE
CubeSats.
3. Orbital correction maneuver relayed via
DSN. Inbound excess velocity via radio
Doppler.
4. Flyby: GPS data from spacecraft to
ground stations. Ground station
measured Doppler shift. Possible SLR
position monitoring.
5. Outbound excess velocity via radio
Doppler. Orbital correction maneuver
relayed via DSN.
6. Repeat flyby or disposal based on
system lifetime.

ConOps A: Repeat tandem flybys of Earth
1. Launch as a secondary payload, highly
inclined.
2. SHERPA second stage provides
hyperbolic excess velocity for FLARE
CubeSats.
3. Orbital correction maneuver relayed via
DSN. Inbound excess velocity via radio
Doppler.
ground stations. Ground station
measured Doppler shift. Possible SLR
position monitoring.
5. Outbound excess velocity via radio
Doppler. Orbital correction maneuver
relayed via DSN.
6. Repeat flyby or disposal based on
system lifetime.
From mid 2015 to mid 2018 further project development will take place. Further pre-
phase A and conceptualization will take place during 2015 to mid 2016. Fabrication,
testing, and assembly will take place from mid 2016 to early 2018.

ConOps B: Mother ship deployment, moon assist
1. Launch as secondary payload to a GTO
orbit.
2. SHERPA delivers CubeSats to moon
sphere of influence.
3. Powered flyby of the moon.
4. SHERPA provides hyperbolic excess
velocity. CubeSats deployed into
tandem hyperbolic flyby trajectories.
Inbound excess velocity calculated
(DSN monitored radio Doppler).
ground station. DSN measured Doppler
shift. SLR tracking possibility.
6. Hyperbolic excess velocity calculated
on outbound leg via radio Doppler.
7. System disposal.

Day in the Life: CubeSat Orientation
• Heliocentric
– Solar panels point toward sun intermittently
– Stable spin (Z-axis)/tumble to distribute heat passively
– MCM maneuver and reaction wheel desaturation
– Small course correction in weeks leading to and days following flyby
• In/out bound flyby
– X-band patch antennas (± Z faces) face towards DSN station of interest
– A slow spin about the Z-axis wouldn’t distort data (preprocessed signal).
– Trajectory profile gathered in intervals
• Closest approach flyby
– SLR reflector (± Z faces) would be pointed towards the relevant station
– GPS signals received, stored and then relayed when appropriate
– (optional) Radio tracking via relevant station (DSN or ESA based on position and slew rate)
– No spin about the Z-axis is preferred due to the slewing necessity at this phase

Baseline Design
CubeSat PBS, orange = primary to mission, yellow = data source, red = in contention.

Baseline Design
INSPIRE cubesat provided for subsystem design heritage (left) and Iris X-Band transponder system (right), courtesy of JPL [33].

Comms Design:
Alternative
JPL designed X band transponder [34].
Design resource for Ling Budget and
comms system characteristics [34]:
• 1 U with 0.5U goal
• ~1 Kg
• 8 W, active with ~3 W goal
• ~>1 m ranging accuracy
• Goal of ~$100k unit cost

INSPRE
configuration
using an X-
Band LMRST
Comms
system [45].

X-Band
LMRST
Comms
Link
Budget
[34].

Radio Aurora eXplorer
(RAX) board assembly
and fully assembled
CubeSat [43]
Antcom L1 GPS patch
Antenna P/N 1.5G15A
Link Budget Example in
LEO.
This report from the
University of Michigan
was intended to assist
future CubeSat missions
in regards to GPS and
link budget [43].
GPS Comms Link Budget and Design

Radio Aurora eXplorer
(RAX) mission GPS
Comms Link Budget for
reference [43].
GPS Comms Link Budget and Design

Baseline Trajectory
CONOPS A

Baseline Trajectory
CONOPS A
Disposal
Date 05/15/2018
ΔV (m/s) Remaining
a (km) 1.1612E+08
e 0.1230
i (deg) 68.567
RAAN (deg) 81.239
w (deg) -52.467

Baseline Trajectory
CONOPS B
Flyby
ΔV (m/s) 0.00
V_inf (km/s) 2.00
a (km) -9.94E+04
e 1.0827
H (km) 1833
DA (deg) 84.6
Vp (km/s) 9.648
δi (deg) 56.4
δo (deg) -28.2
ΔV_inf (mm/s) -1.017

DeltaV Budget
Timeline Event
ConOps A –
SHERPA
ConOps A –
FLARE
ConOps B –
SHERPA
ConOps B –
FLARE
Departure 1.407 0.7163
MCM1 0.050 0.010
MCM2 0.100 0.020 0.010
Flyby1
MCM3 0.020 1.9754
MCM4 0.020 0.010 0.030
Flyby2
Disposal 0.6430 0.040 0.020
TOTAL 2.200 0.100 2.722 0.050
MARGIN 0.400 0.050 -0.122 0.025
AVAILABLE 2.600 0.150 2.600 0.075

Tracking Trade Studies
• Slew Rate
– DSN: 70m dish, 0.25 deg/s
34m dish, 0.4 deg/s
– Estrack: 34m DSA1, 0.4 deg/s
34m DSA2, 1.0 deg/s
– TDRSS: 1.0 deg/s
– Worst Case: 0.35 deg/s
• Visibility
– DSN: Full visibility at > 30,000 km, low
visibility for Earth orbits (few stations)
– Estrack: High visibility for Earth orbits
due to cooperating networks
– TDRSS: Full visibility

• Slew Rate
– DSN: 70m dish, 0.25 deg/s
34m dish, 0.4 deg/s
– Estrack: 34m DSA1, 0.4 deg/s
34m DSA2, 1.0 deg/s
– TDRSS: 1.0 deg/s
• Visibility
– DSN: Full visibility at > 30,000 km, low
visibility for Earth orbits (few stations)
– Estrack: High visibility for Earth orbits
due to cooperating networks
– TDRSS: Full visibility (< 12,000 km)

• Slew Rate

Trade Study: Separation
• Considerations:
– Trackability sets inner bounds.
• Assume single receiver, slew rate to track, then reset for second pass.
• FLARE perigee pass length: ~2597s for 180 degrees.
• Slew time to return to track 2nd satellite: 450s.
– Minimum separation of 3047s = 11651km at Vinf.
– Similitude sets outer bounds.
• Orbit does not depend on planetary synodic periods.
• Important parameter is direction of inclination of equator to the ecliptic, rate of change 0.99 deg./day.
• For small angle, change results in increased MCM to achieve heliocentric transit.
– Increased separation increases deltaV for MCM & separation maneuver.
– Select separation near minimum w/ safety margin, ~6000s = 22,942km.

Trade Study: Satellite Laser Ranging [46,47]
• SLR: Satellite Laser Ranging measures the travel time of light pulses from
a ground station to a spacecraft and back
• Spacecraft must have special reflector attached
• Altitudes from 300-22,000+km
• SLR: Current accuracy on the order of millimeters
• 1-2mm normal point precision
• Ground Stations available in: USA, Hawai’i, Peru, Australia, South Africa,
and Tahiti allowing global coverage
Speaker: Anthony Huet
[46]

Trade Study: Propulsion Subsystem
• Hydrazine Propulsion
– Lots of heritage with spacecraft
– Simpler implementation
– Relatively high thrust
• Electric Propulsion
– Candidate solution for Primary ConOps
– Highest ISP, low thrust but sufficient time for burn
– Low TRL
• Cold Gas Propulsion
– Candidate solution for Secondary ConOps
– Lowest ISP, moderate thrust
– Simple

Element CBE Contingency (10%) Allocated Level 1
1.0 Spacecraft Bus 9189 g
1.1 Propulsion (WET)
MPS-120XL™ CubeSat High-Impulse Adaptable13
3200 g 320 g 3520 g
1.2 ADCS
BCT XACT10
850 g 85 g 935 g
1.3 Communication
Iris Navigation and Telecomm Transponder 400 g 40 g 440 g
1.4 C&DH
Andrews Model 160 High Performance Flight Computer9
70 g 7 g 77 g
1.5 Power
12
FleXible EPS 6 x 12W BCR 139 g 13.9 g 152.9 g
CubeSat Power Distribution Module 61 g 6.1 g 67.1 g
CubeSat Standalone Battery 256 g 25.6 g 281.6 g
6U CubeSat SIDE Solar Panel 290 g 29 g 319 g
6U CubeSat SIDE Solar Panel 290 g 29 g 319 g
3U CubeSat Side Solar Panel 135 g 13.5 g 148.5 g
3U CubeSat Side Solar Panel 135 g 13.5 g 148.5 g
1.6 Structure
6-Unit CubeSat Structure
9
1100 g 110 g 1210 g
1.7 Sensors
FOTON GPS Receiver 400 g 40 g 440 g
1.8 Wiring
15 % Of components not including structure 1027 g 102.729 g 1130 g
2.0 Margin (15% of 1.0) 1378 g
3.0 Total CubeSat Mass 10567 g
Level 2
Master Equipment List (MEL)

• The volume analysis of the actual components used is displayed in the following
graph
Volume Analysis
• 3.85 U, assuming 96x96mm base, using maximum volume components
Type Product Size (mm) Height (mm)
Height Contingency
(10%)
Total Height
(mm)
RW&C BCT XACT10 100x100x50 mm (0.5U) 50 5 55
Sensors None (BCT XACT) None 0 0 0
Radio ISIS VHF/UHF Full Duplex Transceiver9 96x90x15 mm 15 1.5 16.5
GPS SGR-05U - Space GPS Receiver14 70x46x12 mm 12 1.2 13.2
Computer ISIS On Board Computer9 96x90x12.4 mm 12 1.2 13.2
EPS FleXible EPS 6 x 12W BCR12 15.3 mm Height, (1 U base area) 15.3 1.5 16.8
Power CubeSat Power Distribution Module12 91x90.5x25 mm 25 2.5 27.5
Batteries CubeSat Standalone Battery12 95.885x90.17x22.215 mm 22.2 2.2 24.4
Solar Panels 6U CubeSat SIDE Solar Panel12 No size inside cubesat 0 0 0
Propulsion MPS-120XW™ CubeSat High-Impulse Adaptable13 200x100x113.5 mm 113.5 11.35 124.85
Sub-Total Height (mm)
335.24
Margin Height Margin
50.29
Total Height (mm) Volume (U)
385.53 3.86

Power Equipment List (PEL): Nominal Power Usage
Speaker: Anthony Huet or Kyle Chaffin
1.0 Spacecraft Bus 18.073 W
1.1 ADACS
BCT XACT 2.83 W 0.283 W 3.113 W
1.2 Radio
Iris Navigation and Telecomm Transponder 6.4 W 0.64 W 7.04 W
1.3 GPS
FOTON GPS Receiver 1 W 0.1 W 1.1 W
1.4 Flight Computer
Andrews Model 160 High Performance Flight Computer 5 W 0.5 W 5.5 W
1.5 EPS
FleXible EPS 6 x 12W BCR 0.1 W 0.01 W 0.11 W
1.6 Batteries
CubeSat Standalone Battery 0.1 W 0.01 W 0.11 W
1.7 Propulsion
MPS-120XL™ CubeSat High-Impulse Adaptable 1 W 0.1 W 1.1 W
2.0 Margin (15% of 1.0 2.71095 W
3.0 Total Nominal Power Usage 20.784 W
Level 2
Element12 Power Output
6U CubeSat SIDE Solar Panel 18.78 W
3U CubeSat Side Solar Panel 7.3 W
Total Power Output 52.16 W
40% Total Power Output
20.86 W
36.51 W

Speaker: Anthony Huet or Kyle Chaffin
1.1 ADACS
BCT XACT 3 W 0.283 W 3.113 W
1.2 Radio
Iris Navigation and Telecomm Transponder 6 W 0.64 W 7.04 W
1.3 GPS
FOTON GPS Receiver 5 W 0.45 W 4.95 W
1.4 Flight Computer
1.5 EPS
FleXible EPS 6 x 12W BCR 0 W 0.01 W 0.11 W
1.6 Batteries
CubeSat Standalone Battery 0 W 0.01 W 0.11 W
1.7 Propulsion
2.0 Margin (15% of 1.0 4.44345 W
3.0 Total Maximum Power Usage 34.0665 W
Level 2
Power Equipment List (PEL): Maximum Power Usage
Element12 Power Output
Total Power Output 52.16 W
20.86 W
36.51 W

Power Equipment List (PEL): Desaturation Power Usage
Speaker: Anthony Huet or Amritpreet Kang
1.1 ADACS
BCT XACT 2.83 W 0.283 W 3.113 W
1.2 Radio
1.3 Flight Computer
1.4 EPS
1.5 Batteries
1.6 Propulsion
2.0 Margin (15% of 1.0 3.70095 W
3.0 Total Desaturation Power Usage 28.374 W
Level 2

Power Equipment List (PEL): Flyby Power Usage
Speaker: Anthony Huet or Amritpreet Kang
1.1 ADACS
BCT XACT 2.83 W 0.283 W 3.113 W
1.2 Radio
1.3 GPS
FOTON GPS Receiver 4.5 W 0.45 W 4.95 W
1.4 Flight Computer
1.5 EPS
1.6 Batteries
2.0 Margin (15% of 1.0 3.78345 W
3.0 Total Flyby Power Usage 29.0065 W
Level 2

Iris:
Comms Link
Budget
Down Link
Rates for the
INSPIRE
CubeSat [33].

Iris:
Down Link Rates,
courtesy of JPL [34]

Cost Analysis: Components of One 6U CubeSat
Type Product Cost Source
Structure 6-Unit CubeSat Structure $8,242 Directly from cubesatshop.com
ADACS BCT XACT $139,995 Directly from pumpkininc.com
Radio Iris Navigation and Telecomm Transponder $10,000 Estimated from cubesatshop.com
GPS FOTON GPS Receiver $50,000 Directly from Brumbaugh Thesis
Flight Computer Andrews Model 160 High Performance Flight Computer $53,261 Directly from cubesatshop.com
EPS FleXible EPS 6 x 12W BCR $10,550 Directly from clyde-space.com
Power Dist. CubeSat Power Distribution Module $8,450 Directly from clyde-space.com
Batteries CubeSat Standalone Battery $1,800 Directly from clyde-space.com
Solar Panels 6U CubeSat SIDE Solar Panel $14,300 Directly from clyde-space.com
6U CubeSat SIDE Solar Panel $14,300 Directly from clyde-space.com
3U CubeSat Side Solar Panel $6,050 Directly from clyde-space.com
3U CubeSat Side Solar Panel $6,050 Directly from clyde-space.com
Propulsion MPS-120XL™ CubeSat High-Impulse Adaptable $125,000 Estimated from tudelft.nl
Wiring 15 % Of components not including structure $44,799.80 10% of Other Product Costs
Total Cubesat Cost: $492,798

Cost Analysis: Two 6U CubeSats and Operations
WBS Element Input CER ($K FY 15) Cost Driver(s) Input Range
1.1 Spacecraft & Payload $985.60 K Component Cost Analysis
1.2 IA&T $985.60 K $137.00 K Spacecraft Bus Total Cost ($K) 2600-69000 ($K FY 15)
3.0 Program Level $985.60 K $225.70 K Spacecraft Bus Total Cost ($K) 2600-69000 ($K FY 15)
4.0 LOOS $985.60 K $60.12 K Spacecraft Bus Total Cost ($K) 2600-69000 ($K FY 15)
5.0 GSE $985.60 K $65.05 K Spacecraft Bus Total Cost ($K) 2600-69000 ($K FY 15)
Total Cost: $1,473,465.42
1.2 Spacecraft Integration, Assembly, and Test
5.0 Aerospace Ground Equipment
4.0 Flight Support
3.0 Program Level
1.1 Spacecraft & Payload

Risk
• Largest risk from component failure
– Radiation hardened components
– Redundant systems

Mandatory Considerations: External Issues
• Economics, Environmental and Sustainability
– Low total cost, but likely low science return.
– Environmental effects consistent with primary payload.
• Disposal options do not increase orbital debris issues.
• Does not add to primary mission environmental impact
• Ethical, Social and Health/Safety
– Ethically and socially pertinent to improving propagation of near-Earth bodies.
– Hydrazine is a health risk. High TRL mitigates.
• Manufacturability, Political and Global Impact
– Developed using standard, high TRL bus and components. Easily manufactured.
– Flyby altitudes well over LEO, small collision probability during flybys.
– Requires/facilitates intl. cooperation of ground stations/launch facilities.

Critical Issues
• Reevaluate design choice based on an empirical trade study
• Radiation exposure during heliocentric trajectories
• Attitude capabilities for “quiet flyby” scenario
• Thermal requirements
• Tracking ability during flyby
• Comms Link Budget
• JPL anomaly explanations
Element Product Minimum Maximum
Structure 6-Unit CubeSat Structure -40°C 80°C
Flight Computer Andrews Model 160 High Performance Flight Computer -30C 65°C
Power Dist. CubeSat Power Distribution Module -40°C 85°C
Batteries CubeSat Standalone Battery -10°C 50°C
Propulsion MPS-120XL™ CubeSat High-Impulse Adaptable 5°C 50°C
Overall Thermal Limits 5°C 50°C
Operating Temperature

Questions and Comments
Project Manager:
Amritpreet Kang
Systems Engineer:
Graeme Ramsey
Chief Engineer:
Jeffrey Alfaro
Associate Engineers:
Kyle Chaffin
Anthony Huet
Graphic courtesy of NASA
𝐾 =
2ω 𝑒 𝑅 𝑒
𝑐
= 3.099 x 10−6

References
• [1] Michael M. Nieto and John D. Anderson. “Earth flyby anomalies”, Physics Today. Oct 2009.
• [2] Anderson, John D., Campbell, James, K., “Anomalous Orbital Energy Changes Observed during Spacecraft Flybys of Earth”. JPL. March
2008. Web. <http://journals.aps.org/prl/pdf/10.1103/PhysRevLett.100.091102>
• [3] Jason Andrews. “Spaceflight Secondary Payload System (SSPS) and SHERPA Tug - A New Business Model for Secondary and Hosted
Payloads”, Spaceflight, Inc. 26th Annual AIAA/USU Conference on Small Satellites.
• [4] “CANISTERIZED SATELLITE DISPENSER (CSD) DATA SHEET”, Planetary Systems Corporation. 21 Jul 2014.
• [5] “Space Launch Report: Rokot/Strela”, http://www.spacelaunchreport.com/rokot.html#config. 19 Dec 2014.
• [6] Antreasian, P., Guinn, J., “Investigations Into the Unexpected Delta-V Increases During the Earth Gravity Assists of Galileo and NEAR”.
JPL. Web.
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Trade Studies: Doppler and GPS Heritage
• Data Acquisition: Projected Accuracy
– Doppler
• Ionosphere (and solar wind) refractive index is proportional to λ^2 (wavelength squared). [20]
• Prospective Europa Orbiter Mission: estimated X band Doppler shift of 0.1 mm/s [6.7e-13 Allen Deviations in 2 way X-band
Doppler at 60 seconds integration which is equivalent to 4e-13 over 1000s integration]. [20]
• Iris and X/X LMRST projected to attain 0.1 mm/s accuracy, estimate courtesy of JPL.
– GPS real time processing
• CanX-2 mission: OEM4-G2L receiver: 10-100 m position error, 0.1-0.5 m/s velocity error [limited by employed antenna].
[16]
• Radio Aurora Explorer II mission: RAX-2 receiver: 2.9 m, .34 m/s average position and velocity error. [17]
– GPS post processing
• CHAMP mission: JPL developed BlackJack GPS: Bernese 4.3 and 5.0 software packages respectively (post processing)
produced 0.5 mm/s and 0.1 mm/s error. [18]
• Multi-GNSS real time tracking offers ~20 mm/s in accuracy in HEO. With weak signal and offline processing 1.0 mm/s
accuracy is achievable. [15]

Table of steady-state navigation errors [21], for
analysis of expected accuracies. Two perigee
passes were necessary to achieve this level of
steady-state accuracy.

SME-SMAD WBS Element Input CER ($K FY 10) CER ($K FY15) Cost Driver(s) Input Range Standard Error (absolute)
1.1a Spacecraft Bus (alternate) 10.1756 kg $1,725.86 K $1,857.76 K Spacecraft Bus Dry Weight 20-400 kg $3,696.00 K
1.1b Spacecraft Bus $8,862.04 K $9,539.37 K Sum of Spacecraft Bus Elements ($)
1.1.1 Structure 2.42 kg $448.28 K $482.54 K Structure Weight (kg) 5-100 kg $1,097.00 K
1.1.2a Thermal Control 10 kg $905.00 K $974.17 K Min. Thermal Control Weight (kg) 5-12 kg $119.00 K
1.1.2b Thermal Control 24 kg $3,618.20 K $3,894.74 K Max. Thermal Control Weight (kg) 5-12 kg $119.00 K
1.1.3 ADCS 1.87 kg $1,890.91 K $2,035.44 K ADCS Weight (kg) 1-25 kg $1,113.00 K
1.1.4 EPS 0.3058 kg $1,490.67 K $1,604.60 K EPS Weight (kg) 7-70 kg $910.00 K
1.1.5 Propulsion (Reaction Control) 10.1756 kg $144.93 K $156.01 K Bus Dry Weight (kg) 20-400 kg $310.00 K
1.1.6a TT&C 1.76 kg $605.05 K $651.30 K TT&C Weight (kg) 3-30 kg $629.00 K
1.1.6b C&DH 0.154 kg $664.00 K $714.75 K C&DH Wight (kg) 3-30 kg $854.00 K
1.2 Payload $8,862.04 K $3,544.82 K $3,815.75 K Spacecraft Bus Total Cost ($K) 2600-69000 ($K FY 10)
1.3 IA&T $8,862.04 K $1,231.82 K $1,325.97 K Spacecraft Bus Total Cost ($K) 2600-69000 ($K FY 10)
4.0 Program Level $8,862.04 K $2,029.41 K $2,184.52 K Spacecraft Bus Total Cost ($K) 2600-69000 ($K FY 10)
5.0 LOOS $8,862.04 K $540.58 K $581.90 K Spacecraft Bus Total Cost ($K) 2600-69000 ($K FY 10)
6.0 GSE $8,862.04 K $584.89 K $629.60 K Spacecraft Bus Total Cost ($K) 2600-69000 ($K FY 10)
Total Cost: $18,077,105.98
1.1 Spacecraft
6.0 Aerospace Ground Equipment
5.0 Flight Support
4.0 Program Level
1.3 Spacecraft Integration, Assembly, and Test
1.2 Payload
Small Spacecraft Cost Model (SSCM):
not accurate, shows current cost model inadequacies
Background | Mission Statements | Requirements | Constraints | ConOps | Baseline | Trade Studies | Design Selection xx

Trade Study: Velocity Data Accuracy Comparison
• Data Acquisition: Position and Velocity
– GNSS/GPS
• Multitude of Ground Stations (NEN)
• L1/L2 (dual band or more)
• Will serve primarily as complimentary data, and near periapse coverage
• Cross link ranging option
– Radio Doppler Monitoring
• DSN
• Dual frequency
• Slew rate of DSN limits coverage near periapse
– Deployable dish vs. patch antenna
• Earth SOI is at about .006 AU
• Noise considerations for data acquisition
– X- vs. S- vs. Ka-band
• Noise mitigation
• Velocity accuracy over our range (~.006 AU)
– Relative position system
• Requires inordinate power and pointing precision
– SLR
• Closest approach coverage
• No onboard power requirement
• Coverage provided by eight particular ground stations
Range (AU)
Data
Rate
(bps)
L1
.01AU
Moon
.0026A
U
Figure courtesy of NASA JPL [19]

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