FLARE report pre-finalized version

FLyby Anomaly Research Endeavor
FLARE Final Report
Graeme Ramsey, Jeffrey Alfaro, Amritpreet Kang, Kyle Chaffin, and Anthony Huet
May 08, 2015
ASE 374L Spacecraft/Mission Design: Dr. Fowler
The University of Texas at Austin
In conjunction with JPL: Travis Imken and Damon Landau
Spring 2015
*point mass orbital mechanics, 2D flyby visual

Table of Contents
[RED=needs corrected (mostly from Fowler and Travis comments),
BLUE= No commentary not corrected,
GREEN=New not finished,
black=corrected]
****(optional) after everything is finalized, If someone wants to hyperlink the table of contents
to the sections, that would be professional. hyperlinking the figures/tables would also be pro,
less important though.
Executive Summary
1.0 Introduction
1.1 Heritage
1.1.1 Phenomenological Formula (NEW)
1.2 Mission Motivations
1.3 Unconfirmed Explanations of the Flyby Anomaly
1.3.1 Dark Matter Encircling the Earth
1.3.2 Modifications in Inertia
1.3.3 Special Relativity [needs anisotropy of light discussion]
1.3.4 Lorentz Accelerations
1.3.5 JUNO Findings: Higher Order Gravity Terms [emphasis: JUNO and
anisotropy as most likely candidates]
1.4 Mission Constraints and Assumptions [see last paragraph Jeff]
1.5 ConOps [I thought it was good to introduce this to our readers early. it could go
in section 4 though…thoughts?]
1.5.1 Primary ConOps [MCM approx. needed]
1.5.2 Secondary ConOps [GTO launch alternative needed]
1.5.3 Launch Details (NEW) [see Appendix I for relevant figures, that
should save you some time/effort]
1.5.4 Day in the Life of FLARE (NEW)[details needed:dishes,slew rates,etc]
2.0 Driving Statements and Requirements
2.1 Scope
2.1.1 Need
2.1.2 Goal
2.1.3 Objectives
2.1.4 Mission
2.1.5 System Constraints
2.1.6 Assumptions
2.1.7 Authority and Responsibility

2.2 Primary Requirements
2.2.1 Mission Requirements [slew rate sat & DSN needed]
2.2.2 System Requirements
2.2.3 Requirements Traceability Matrix (NEW, needs discussion)
3.0 System DesignDevelopment
3.1 Design Alternatives Development
3.1.1 Preliminary ConOps 1
3.2 System and Subsystems Allocation
3.3 Design Heritage
3.3.1 INSPIRE CubeSat
3.3.2 X/X-band LMRST
3.3.3 Iris X-band Transponder
3.3.4 GPS/GNSS Receivers [addition figure, discussion needed]
3.3.5 Satellite Laser Ranging (NEW)
3.4 Trade Study Summary and Results
3.4.1 Data Acquisition Systems [slew rate discussion, and Amrit GPS resources]
3.4.2 Launch Vehicle
3.4.3 Trajectory [discussion, velocity triangle and departure depiction needed]
3.4.4 Primary (A) vs. Secondary (B) ConOps [empirical trade study needed,
cost, risk, etc needed]
3.4.5 Propulsion
3.4.6 Desired Component Characteristics (NEW)
3.4.7 Prospective Modeling Analysis (NEW)
3.5 Critical Parameters [details needed: DSN coverage/usage, Amrit component
resources, CAD analysis discussion, fix table, etc.]
3.6 Midterm Design Refinement
3.6.1 JPL Midterm Mission Design Presentation Feedback
3.6.2 Launch Vehicle and Launch Trajectory Details [details needed]
3.6.3 Burn at Earth SOI Calculations [details needed]
3.6.4 Subsystem Component Choices [details needed]
3.6.5 CAD Model for Analysis
3.6.6 Final Flyby Maneuver and System Disposal (NEW)
4.0 System Design
4.1 Baseline Designs
4.1.1 Primary ConOps Baseline Trajectory
4.1.2 Primary/Secondary ConOps Evaluation (NEW)
4.2 Design Choice

4.2.1 System and Subsystem Overview [significant updates needed]
[NO FEEDBACK for below sections in blue]
4.2.2 Master Equipment List (MEL)
4.2.3 Equipment Volume Allocation List (EVAL)
4.2.4 Power Equipment List (PEL)
4.2.5 Comms Link Budget and EbNo Analysis (NEW)
4.3 Mission Timeline and Schedule (NEWish)
4.4 Cost Analysis (NEW)
4.5 Risk Analysis (NEW)
4.6 Mandatory Considerations (allNEW)
4.6.1 Economics, Environmental and Sustainability Issues
4.6.2 Ethical, Social and Health/Safety Issues
4.6.3 Manufacturability, Political and Global Impact Issues
5.0 Summary and Conclusions [corrected...but will need final update]
6.0 DesignCritique (all NEW)
6.1 Strengths
6.2 Weaknesses
6.3 Confidence
6.4 Alternatives
7.0 References
8.0 Appendices
Appendix I: Primary Resources Reference Information [add anything important]
Appendix II: FLARE Team Management[updates needed>>tables and contribution
statements]
Appendix III: Subsystem Requirements
Appendix IV: JPL Feedback

List of Tables
Table 1: Flyby orbital parameters of heritage missions[add JUNO data]
Table 2: Heritage Missions Navigation[model order row needed, also can you make
it the same format as other tables>>Excel]
Table 3: Radio band comparison
Table 4: Design selection criteria
Table 5: Tentative ConOps selection
Table 6: Thermal requirements[needs updated to current components]
Table 7: Baseline trajectory data, departure and heliocentric
Table 8: Baseline trajectory data, flybys
Table 9: MEL
Table 10: EVAL
Table 11: PEL[MEL/PEL/EVAL need cleaned up as previously mentioned, e.g.
contingency is part of total, same format, readable, etc.]
with Figure 15 to identify components, y’all can size down all the extra subsystem description
row labels...this will make it easier to read. Order them the same as Figure 15 gets ordered.
List of Figures
Figure 1: Magnitude of Potential Error Sources
Figure 1: Primary ConOps
Figure 2: Secondary ConOps
Figure 3: CSD dispenser deployment setups
Figure 4: Sherpa on payload section
Figure 5: FLARE Primary ConOps PBS
Figure 6: INSPIRE cubesat
Figure 7: JPL LMRST
Figure 8: Iris X-band Transponder
Figure 9: BlackJack GPS receiver
Figure 10: Radio Aurora eXplorer
Figure 10: Radio band comparison
Figure 11: Launch system analysis
Figure 12: Preliminary CAD model [unless you put actual components into the model
scratch this figure… discussion of the uses of a CAD model is definitely relevant though.]
Figure 13: Baseline trajectory, departure and heliocentric
Figure 14: Baseline trajectory, flybys
Figure 15: Component Selection PBS [needed]
Figure 16: Timeline for primary ConOps

Acronyms and Symbols
~ Approximately
> Greater than
a Semimajor axis
e Eccentricity
i Inclination
H Altitude of periapsis
φ Geocentric Latitude
λ geocentric longitude
Vf Inertial spacecraft velocity at closest approach
V_inf Hyperbolic excess velocity
ΔV_inf Anomalous change in hyperbolic excess velocity
DA Deflection angle
αi Right ascension of the incoming osculating asymptotic velocity vector
δi Inbound declination
αo Right ascension of the outgoing osculating asymptotic velocity vector
δo Outbound declination
ADCS: Attitude Determination and Control System
AU: “Astronomical Unit”, Earth’s approximate distance from the Sun
ConOps: Concept of Operations
CSD: Capsulized Satellite Dispenser
DSN: Deep Space Network
DV: “Delta-V”, a propulsive maneuver resulting in velocity change
EELV: Evolved Expendable Launch Vehicle
EM: Earth to Moon
EPS: Electrical Power System
EVAL: Equipement Volume Allocation List
EVE: Earth Venus Earth, order of flybys on trajectory
FLARE: Flyby Anomaly Research Endeavor
FOTON:
GNSS: Global Navigation Satellite System
GPS: Global Positioning System
GN&C: Guidance Navigation and Control
HEO: High Earth Orbit
JPL: Jet Propulsion Laboratory
J#: Gravity term of denoted order (#)
LEO: Low Earth Orbit
LMRST: Low Mass Radio Science Transponder
ME: Moon to Earth

MEL: Master Equipment List
NEN: Near Earth Network
PBS: Product Breakdown Structure
PEL: Power Equipment List
RAAN: Right Ascension of the Ascending Node
RAX: Radio Aurora eXplorer
S/C: Spacecraft
SOI: Sphere of Influence
SLR: Satellite Laser Ranging
SSPS: Spaceflight Secondary Payload System
TBR: To Be Resolved
TPS: Thermal Protectant System
TRL: Technology Readiness Level
wrt: With Respect To

Executive Summary
Planetary flybys have been in use since Mariner 2 flew by Venus in 1962. Team FLARE
(FLyby Anomaly Research Endeavor) at 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, Rosetta and most recently JUNO during flybys of Earth. The anomaly takes
the form of an unaccounted for change in energy/velocity which has observed taking place near
periapse of hyperbolic Earth flybys. The anomaly’s magnitude is linked to the relative velocity
of the spacecraft and inbound/outbound declinations. Although the anomaly has only been
realized and measured in Earth flybys, it is likely present in captured orbits as well, just much
less notable in magnitude. This project has merits in regards to refining our current
understanding of (planetary level) physics and particularly the modeling of near Earth or Earth
rendezvousing objects (e.g. asteroids). It could also result in more precise trajectory modeling
and tailored use of the “anomalous” velocity change to suit particular mission trajectories
(especially regarding Jupiter [or Sun] flybys which would produce the largest anomaly in our
solar system).
The recorded velocity anomalies vary by as much as 13.5 mm/s from modeled values.
These anomalies fit a phenomenological formula which relates the velocity discrepancy to excess
velocity, change in declination and a constant scaling factor involving the ratio of Earth’s
angular velocity times its radius, to the speed of light. The formula isn’t precise and only fits
anomalies where closest approach took place under 2000 km. Many possible causes have been
conjectured, accounted for, or proved innocent (like atmospheric drag and J2 effects). Initially a
thorough investigation of the navigation software and mathematical models used for navigation
by JPL uncovered no hint of the culprit. Early conjectured sources of the anomaly include
unaccounted for relativistic effects, high order gravity terms stacking, atmospheric drag, tidal
effects, Lorentz acceleration, inertial effects or even dark matter. Further investigation by JPL
uncovered two most likely sources of the anomaly, modeling errors that might take the form of
high order gravity terms or, alternatively, the 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. In accomplishing that goal we intend to use high technology
readiness level (TRL) components and redundant/complementary platforms for tandem data
retrieval. The primary Concept of Operations (ConOps) incorporates a heliocentric trajectory
where an unpowered Earth flyby should be executed on an alternating six monthly and yearly
basis (approximately). A secondary ConOps incorporates a powered flyby of the moon followed
by a single unpowered flyby event (meaning multiple deployed-satellite trajectories on one
flyby) of Earth. The hope is to get at least 4 more data points to compliment the current data on
the anomaly. To demonstrate repeatability, the satellites will fly in pairs on tandem trajectories.
To reflect the project’s tentative budget of $5mil excluding launch associated costs, the satellite
design will be limited to 6u cubesats. It was assumed (in regards to the primary ConOps) that
our satellites would have a lifetime of at least 2 years, and that launches as a secondary payload
to an inclined (~60 deg with respect to Earth’s equator), highly elliptic (~0.74) and suitably

elevated (apogee altitude ~ 40,000 km) parking orbit would be within our budget. Other
assumptions are a 10-15% mass/volume/power contingency and 40% sunlight exposure for static
solar arrays.
The primary considerations for the FLARE mission are: a) design a cubesat system
capable of facilitating velocity measurements accurate to the order of 0.1 mm/s, b) perform
multiple Earth flybys with regards to the phenomenological formula, c) if possible, gather data in
a manner to help characterize the anomaly. The data acquisition system trade study in regards to
accuracy of velocity measurements is paramount for this mission. The anomaly is on the order
of mm/s and must be observable by the space and Earth bound systems. The Earth based
systems include the Global Positioning System (GPS) and radio (X/S-Band) doppler monitoring
via ground stations (Near Earth Network [for GPS] and Deep Space Network [for radio]) with
post-processing, and possibly Satellite Laser Ranging as a compliment or substitute for GPS.
The trajectory coupled with primary propulsion system trade studies have broad trajectory design
ramifications as well as redistributing the mass/volume and power budgets. High order gravity
terms (modeling up to >J120) have been conjectured as the most probable cause of the anomaly.
A trade study on this subject to apply new gravity models, acquired from missions like GRACE
(Gravity Recovery and Climate Experiment), to our heritage missions could supply evidence that
the source of the anomaly is a modeling error. Contained in the overall report are both technical
and managerial designs(primarily in the appendix).

1.0 Introduction
Gravity assists for spacecraft are well understood maneuvers that have been used for
decades to reach remote locations in the solar system, and, in the case of the Voyager probes, to
escape the solar system. In these hyperbolic flybys the passing spacecraft exchanges heliocentric
orbital energy with the planet, which results in a significant heliocentric velocity vector change
for the spacecraft. The purpose of these flybys is twofold. Current spaceflight technology does
not provide enough DV for spacecraft to reach some distant destinations in the solar system, and
the velocity increase can also significantly reduce travel time, capable of reducing mission travel
time by years.
The exact position, angle, and velocity changes experienced by the spacecraft are
calculated to great precision. High accuracy knowledge of the solar system and physics allows
the velocity profile to be modeled to greater degree of accuracy than millimeter per second.
Despite this, during some flybys of the Earth the velocity boost that the spacecraft received was
different than what was initially modeled. The difference was only on the order of millimeters
per second, but remained significant nonetheless. These values were calculated to high precision
using Doppler residuals from the spacecrafts' telemetry data. There does appear to be an
association between asymmetric incoming and outgoing declination angles about the equator and
higher velocity discrepancies, possibly suggesting the anomaly may be the result of some effect
derived from of the rotation of the Earth, but to date there is not a sufficient explanation for the
cause of this occurrence and thus it remains an anomaly. One potential solution posits that the
anomaly can be resolved by using a higher order gravity field of the Earth than was used for the
initial DV calculations. The proposed mission would be the first of its kind to be launched solely
to investigate this anomaly.
1.1 Heritage
While no heritage missions have been dedicated entirely to the study of flyby anomalies,
flyby anomalies have been measured indirectly as part of other missions, such as the ones
mentioned in Figure 1, namely Galileo, NEAR, Cassini, Rosetta, and Messenger. From these
missions, we gather information pertaining to where flyby anomalies occur as spacecraft perform
an Earth flyby, by which we can attempt to reproduce such flyby anomalies in an effort to
determine their existence. For each of these missions, we have data for important orbital
parameters such as height, geocentric longitude and latitude, inertial spacecraft velocity at closest
approach, osculating hyperbolic excess velocity, the deflection angle between incoming and
outgoing asymptotic velocity vectors, the inclination of the orbital plane on the Earth’s equator,
the right ascension and declination of the incoming and outgoing osculating asymptotic velocity
vectors, and an estimate of the total mass of the spacecraft during the encounter [6].

Table 1: Flyby orbital parameters of heritage missions [2]
[Add JUNO characteristics, at least the highlighted ones if possible]
Information pertaining to the communication subsystem of the flyby anomaly heritage
missions are presented in Table 1, which presents the manner in which velocity changes were
measured in heritage missions as well as the means of communicating said changes. As the data
in Table 1 reveals, the velocity measurements of the heritage missions were precise up to 1/100
of a millimeter per second. The missions further display commonality in that they all used X-
band frequency to transmit data, and the velocity in each of the missions was measured by
doppler shift.

Galileo NEAR Cassini Rosetta MESSENG
ER
Juno
Velocity
Measurement
Doppler shift Doppler
shift
Doppler
shift
Doppler
shift
Doppler
shift
Doppler
shift
Band S and X X X S and X X X and Ka
Antenna
Diameter or
Dimensions
(m)
4.6 1.5 4 2.2 0.28 x 0.81 2.5
Velocity
Precision
(mm/s)
0.01 0.01 0.01 0.01 0.01 0.01
Radio Power
(W)
23 5 13.96 20 10 25
Table 2: Heritage Missions Navigation [24-26, 26].
1.1.1 Phenomenological Formula
1.2 MissionMotivations
The FLARE mission is devoted to proving the existence of a physical phenomenon
related to the energy associated with planetary flybys being dissimilar to current orbital
trajectory models. Pertaining to the data gathered during closest approach, this would fill in the
gap left by most of the heritage missions. In the process of gathering more data points to prove
the anomaly’s existence, providing coverage during closest approach could serve to help
characterize the anomaly to a more proficient degree and consequently refine the
phenomenological formula associated with the anomaly.
This project has merits in regards to refining our current understanding of planetary level
physics. FLARE could also result in more precise trajectory modeling and tailored use of the
“anomalous” velocity change to suit particular mission trajectories, thereby saving investment in
fuel mass and mass associated costs. This mission seeks to gather data and understanding in
regards to the inner workings of large scale physics, and in doing so benefit the science
community and aerospace industry as a whole. Of particular relevance, the modeling of near
Earth or Earth rendezvousing objects (e.g. asteroids) would be improved by this mission.
Although the anomaly itself is small, the effect of a small perturbation can become large over
vast distances (e.g. the Voyager satellite velocity magnitude discrepancy).
Other benefits from this project include further advancing the state of the art in regards to
the usage of cubesats in (semi-) deep space missions. It would also serve to further demonstrate

and/or refine emerging cubesat technologies and techniques in regards to navigation in
heliocentric space (trajectory, attitude, radiation mitigation, etc) and cooperative systems.
Secondary payload capabilities would be tested and refined via use of a Spaceflight Secondary
Payload System (SSPS) and a standardized Capsulized Satellite Dispenser (CSD) layout. The
reuse of the SSPS(for means other than as an exit assist vehicle) in conjunction with the cubesats
could serve to advance the state of the art of cooperative (constellation-like) systems, with
deployed cubesats in a semi-static formation and use of a “mothership”.
1.3 Unconfirmed Explanations of the Flyby Anomaly
Several theories have been proposed as explanations for the existence of flyby anomalies,
but as the following subsections should make clear, more data is needed to determine the
existence and nature of flyby anomalies. Figure 1 below depicts the magnitudes of some
perturbations associated with satellites in space. This figure serve to give a point of reference for
what perturbations might resemble the anomaly in magnitude (~10^-5 to ~10^-6).
Figure 1: Magnitude of Potential Error Sources courtesy of a Portuguese mission proposal
regarding examination of the anomaly using GNSS [39].
1.3.1 Dark Matter Encircling the Earth
As an explanation for the existence of flyby anomalies, dark matter encircling the Earth
was offered [28]. It was thought that flyby anomalies could result from the scattering of
spacecraft nucleons due to dark matter particles orbiting Earth. Velocity decreases would be due
to elastic scattering, and velocity increases would arise from exothermic inelastic scattering [28].
However, this theory predicted a large change in change in Juno’s hyperbolic excess velocity of

11.6mm/s [28], but no anomalous change in hyperbolic excess velocity was observed in Juno’s
flyby of Earth [29]. Clearly, another explanation is desired, and FLARE should go a long way in
providing data for the study of flyby anomalies.
1.3.2 Modifications in Inertia
It was attempted to explain the existence of flyby anomalies by a modification of inertia
[30], with the conclusion that a model of modified inertia which used a Hubble-scale Casmir
effect could predict anomalous changes in orbital energy on the order of magnitude of the flyby
anomalies minus NEAR [30]. However, this explanation lacks experimental testing and
empirical data, and it still leaves NEAR as an anomaly among anomalies, unable to accurately
predict its large change in hyperbolic excess velocity.
1.3.3 Special Relativity[where is the anisotropy of the speed of light discussion?]
Special relativity was offered as an explanation for spacecraft flyby anomalies [31]. It
was found that the special relativity time dilation and Doppler shift, along with the addition of
velocities to account for Earth’s rotation pose a solution to an empirical formula for flyby
anomalies [31]. It was thus concluded that spacecraft flybys of heavenly bodies may be viewed
as a new test of special relativity which has proven to be successful near Earth [31]. However,
empirical formulas necessitate empirical data, so with the help of FLARE, more measurements
of the flyby anomaly must be made for an empirical formula to be satisfied by sufficient
empirical data.
1.3.4 Lorentz Accelerations
It was thought that Lorentz accelerations associated with electrostatic charging could
account for the existence of flyby anomalies [32]. However, an algorithm based on this theory
could not converge on a solution that fully reproduces the anomalous error in all six orbital
states, so Lorentz accelerations pose an unlikely explanation for the existence of flyby anomalies
[32]. Once again, more data is needed.
1.3.5 JUNO Findings: Higher Order Gravity Terms
On October 9, 2013, the JUNO spacecraft flew by earth with relatively high expected
changes in orbital energy at or near perigee. For instance, Adler’s dark-scattering model for
predicated anomalous changes in orbital energy in earth flybys predicted a change in hyperbolic
excess velocity of 11.6 mm/s [28], while Antreasian and Guinn’s model predicted a change of 7
mm/s [36]. However, no anomalous velocity change was observed at or near perigee [36]. As a
possible explanation, it was noted that 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]. Other possible sources of perturbation such as the three-sigma error in

Earth’s GM and variations in J2 that would not necessarily be well known in a predictive sense
were considered and discredited as explanations, being incapable of reaching a level of
perturbation that would be easily detected in real-time monitoring [36].
However, there is a potential that cumulative effects of high order gravity terms could
produce a perturbation on the order of magnitude seen in the flyby anomaly, mm/s [36]. Such
higher order terms were used in the trajectory prediction of JUNO’s flyby. The trajectory
produced was accurate to the extent that no flyby anomaly was detected. However, this does not
prove that the cause of the difference between JUNO’s experience and the previously flybys
were due to the trajectory prediction using higher order gravity terms. A simulation of the
previous 6 flybys using very high order terms, up to J100, would provide better evidence of
whether the higher order terms can account for the anomaly.
1.4 MissionConstraints and Assumptions
•The flybys must take place around Earth in order to achieve the required velocity
measurement accuracy.
In order to calculate the velocity of a spacecraft to the accuracy necessary to identify the
proposed hyperbolic flyby anomaly, earthbound installations such as the DSN and NEN are
essential. The available technologies and techniques by which to calculate velocity
measurements decrease in accuracy at increasing distance from Earth. These technologies
include radio doppler analysis which requires use of the DSN, GPS which requires access to the
GNSS and the NEN which are much more limited by range (from Earth) than DSN and
potentially SLR which requires access to earthbound laser facilities.
•Flyby characteristics must coincide with phenomenological formula.
The phenomenological formula developed by JPL, which fits the observed anomaly data,
is as follows:
, [1].
From observation of the variables involved, it becomes apparent that in order to produce a viably
measurable anomaly, a large difference in the cosines of in/outbound declinations (>~0.3) and
large hyperbolic excess velocity (>~1 km/s) appear to be required (corresponding to an anomaly
on the order of mm/s).
•Mission budget: $5mil before launch associated costs.
In order to maximize mission viability it is important to be as efficient as possible with
the space-bound system’s mass and pre-launch costs. An estimate of $5mil prior to launch
associated costs, provided by JPL’s Travis Imken, serves to guide the scope of the FLARE
mission. Detailed in 2.1.5 System Constraints, are launch system budgetary considerations.
Approximate Launch Vehicle (LV) and SSPS costs are expanded on in the Cost section (4.4).

•Launch window and parking orbit/exit trajectory characteristics (Primary ConOps).
In order to achieve a heliocentric trajectory that results in properly constrained flybys, the
hyperbolic excess velocity (V_inf) upon Earth departure must be similar to the V_inf required at
the first flyby. Furthermore, the direction of the Earth departure combined with V_inf must be
such that the spacecraft’s initial heliocentric trajectory is very close to that of the Earth with
exception of an inclination change. This constraint should serve to accommodate our baseline
trajectory (detailed in section 4.1.1). This means that, with respect to the ecliptic place, the
spacecraft must leave the Earth’s SOI in the ecliptic z-direction and slightly against the direction
of Earth’s revolution about the sun. This will help achieve a similar orbit period and allow
rendezvous for the first flyby in ~6 months.
To reduce the fuel consumption needed to achieve the necessary departure trajectory, the
initial parking orbit and thus launch trajectory should be highly eccentric and inclined. The
RAAN of the parking orbit or launch must also match the date of departure such that a DV at
perigee places the spacecraft on the proper trajectory, if fuel mass is to be optimized.
[Jeff, can you expand upon the details of our departure trajectory here?]
1.5 ConOps
To provide a flash forward to our current project direction our midterm mission
procedure is provided here in section 1 rather than section 4. It should be referenced as an
introduction to section 4 as well. The development of these midterm ConOps will be outlined
(among other items) in section 3.
1.5.1 Primary ConOps
The primary ConOps (depicted in Figure 1) chosen from several candidates, consists of
tandem hyperbolic flybys of earth by a cubesat pair and heliocentric trajectories of 6 months
alternating with 1 year between flybys. These cubesats will be capable of having their velocity
profile measured to 0.1 mm/s precision while in Earth’s influence, in order to detect and analyze
the anomaly. The SSPS may also function as an additional velocity profile upon flyby. This
ConOps is projected to allow 2 flyby events in 18 months , which will provide 4 data points
demonstrating repeatability from the cubesats and 1 additional data point from the SSPS. (see
section 1.5.3 for launch and deployment details)

Figure 1: Primary ConOps depiction.
1. Launch as a secondary payload, highly inclined.
Our baseline trajectory assumes launch trajectory characteristics of an inclination of
roughly 60 deg and an eccentricity over 0.7. The date for launch would be set for ~2018 if the
project immediately is adopted by NASA or JPL at the conclusion of our study. We modeled the
situation as departing from a molniya type parking (a=~26000 km, e=0.74, i=63 deg) orbit.
Once the launch vehicle deploys its primary payload, the SHERPA 2200 could immediately
deploy and begin the exit trajectory maneuvers if the launch was nicely matched up with our
baseline trajectory. A parking orbit will most likely be necessary due to the fact the launch
trajectory will facilitate the primary payload. In this scenario SHERPA will deploy after the
primary payload perform small orientation maneuvers to align its orbit in preparation for the
baseline exit trajectory. The primary exit DV maneuver will take place at periapse of the parking
orbit.
2. SHERPA second stage provides hyperbolic excess velocity for FLARE CubeSats.
In performing the above mentioned exit trajectory maneuver, the SSPS will provide at
least 1 km/s of excess velocity to the system. If SHERPA can retain ~100 m/s of DV capability,
it can also serve as a data acquisition system to complement the paired cubesats. At this stage
SHERPA and docked cubesats will traverse a heliocentric trajectory on an inclined orbital plane
to the ecliptic. Autonomous attitude adjustments and system management/testing will take place
on each heliocentric trajectory. The first rendezvous with Earth will take place after 180 degs of
orbit (~6 months). Prior to entering Earth’s SOI the cubesats will be deployed and set into their
tandem flyby trajectory.

3. Orbital correction maneuver relayed via DSN. Inbound excess velocity via Doppler.
As mentioned above the approach maneuvers will be relayed via the DSN and should
take place prior to entering Earth’s SOI. Trajectory modeling will have taken place before the
maneuver commands. These maneuvers include reaction wheel desaturation after attitude
stabilization and trajectory corrections to ensure the proper pared flybys and recurrent flyby
trajectory. Upon entering Earth’s SOI the system will go quiet (e.g. no DV), the inbound excess
velocity will be calculated by analyzing radio Doppler effects via DSN. The inbound velocity
profile will be recorded using DSN and the same radio Doppler analysis, by gathering trajectory
information at intervals(DSN detail needed) upon approach.
4. Flyby: GPS/SLR signals from spacecraft to ground stations. NEN monitoring of (position
and) velocity during closest approach. Alternatively ESA ground station monitoring of radio and
radio Doppler for trajectory analysis.
At the closest approach phase, the DSN radio Doppler velocity profile will cut off due to
the limited slew rate of the DSN dishes (ESA stations may be a viable option for closest
approach). Prior to that point GPS (and/or SLR) will begin monitoring the velocity (and less
vital, the position) profile. This should provide sufficiently accurate velocity data throughout
closest approach.
5. Outbound excess velocity via Doppler. Orbital correction maneuver relayed via DSN.
Once the satellites have left closest approach, the DSN will be able to monitor Doppler
data again. Velocity data will be gathered until after the satellites have exited Earth’s SOI. At
this point (done collecting data for post-processing) the s/c will no longer by “quiet” in that they
may desaturate the reaction wheels and perform maneuvers. Furthermore, once the satellites
post-flyby trajectories have been modeled, a trajectory correction maneuver will be necessary to
set up the next flyby.
6. Repeat flyby or disposal based on system lifetime.
Repeat flybys are limited by the lifetime of critical subsystems. The system lifetime
hinges upon subsystems/components surviving the radiation of space at ~1 AU from the Sun
along with propulsion capabilities in reference to essential trajectory corrections and attitude
device desaturation. The propellent system has a lifetime of (DV capable/MCM) flybys with a
10% contingency considered and only approximate MCMs from heritage data[Jeff MCM
resouce?]. At a point suitable close to the system’s end of life, a final maneuver will be required
to facilitate the systems’ disposal. Disposal can be as easy as redirecting the CubeSats into
Earth’s atmosphere to burn up.
1.5.2 Secondary ConOps
The secondary ConOps (depicted in Figure 2) chosen from several candidates, consists of
tandem hyperbolic flybys of earth by a cubesat pairs after a powered flyby of the moon. These
cubesats will be capable of having their velocity profile measured to mm/s precision while in
Earth’s influence, and by that standard capable of observing the anomaly. The SSPS may also

function as an additional velocity profile upon flyby. This ConOps is projected to allow 1 flyby
event in 1 month, which will provide 4 data points demonstrating repeatability from the cubesats
and 1 additional data points from the SSPS. Theoretically this flyby event could be tailored to
allow another flyby within ~1 year providing an additional 5 data points.
Figure 2: Secondary ConOps depiction.
1. Launch as secondary payload.
A near equatorial launch into a high eccentricity (~0.7) and semimajor axis (~26000 km)
parking orbit is required for this ConOps. The date for launch would be set for ~2018 if the
project immediately is adopted by NASA or JPL at the conclusion of our study. We have not
modeled the situation as of yet, so the finer points of the Earth to Moon (EM) trajectory are TBR.
2. SHERPA second stage delivers FLARE CubeSats to moon sphere of influence.
Once SHERPA 2200 deploys, it will enter a parking orbit and outgas systems to negate
that perturbation during the flyby and considering that launch trajectory will facilitate the
primary payload, a parking orbit will allow the EM trajectory to be aligned. In this scenario
SHERPA will deploy after the primary payload, perform small orientation maneuvers to align
the SSPS orbit in preparation for the EM exit trajectory and perform a burn to enter the Moon’s
SOI. The primary exit DV maneuver will take place at periapse of the parking orbit. (GTO
alternative launch discussion)
3. Powered flyby of the moon.

SHERPA will make use of a powered flyby of the moon to swing around in an effort to
set up an semi-unpowered (semi- due to the fact that this system is in Earth’s influence when it
provides a large DV maneuver) ME flyby trajectory. Details TBR.
4.SHERPA provides hyperbolic excess velocity. CubeSats deployed into tandem hyperbolic
flyby trajectories. Excess velocity calculated (DSN-Doppler).
Upon departure from the Moon, the SSPS will spend the entirety of its DV capabilities in
an effort to maximize the hyperbolic excess velocity, and thus measurable anomaly. Once this
maneuver is complete, the cubesats (4-6) will be deployed and oriented to their tandem flyby
trajectories. At this point radio Doppler measurements will be able to start building the
“unpowered” trajectory profile.
5. Flyby: GPS signals from spacecraft to ground station. DSN measured Doppler shift.
The trajectory upon closest approach can be monitored by GPS and the higher altitude
approach/departure trajectory profile will be built primarily from radio Doppler analysis. This
flyby should provide 4 data points regarding the anomaly demonstrating repeatability (2 tandem
cubesats pairs) and 1 additional data point including the SHERPA.
6. System disposal (possible reuse).
Depending on the CubeSats’ capabilities, either system disposal or reuse would be in
order. This ConOps could borrow the baseline trajectory from the Primary ConOps to set up
repeat flybys. However it is more likely that this ConOps will err on the more affordable side.
And by that standard, the cubesats will not be rad hardened, will have minimal propulsion
capabilities, and will have an expected lifetime of months rather than years.
1.5.3 Launch and Deployment Details
1.5.4 Day in the Life of FLARE
In reference to the Primary ConOps, three primary phases of behavior exist. These main
sets of behavior are described as: 1) heliocentric phase, 2) in/out-bound flyby phase, and 3)
closest approach flyby phase.
The heliocentric phase attitude will be such that the CubeSats deployed solar panels are
pointing toward the sun. Minimizing sensitive component radiation exposure is important during
this phase. Radiation shielding would be strategically placed to protect sensitive components in
this particular attitude. During the 6 months to a year that the Cubesats are in heliocentric space
between flybys, the systems will perform maneuvers to desaturate reaction wheels to maintain
proper attitude. These attitude maneuvers should be tied into mid-course maneuvers (used to
optimize the trajectory) in order to make dual use of the burn. In the weeks leading to and days
following the flyby phase the CubeSats’ trajectory will be analyzed to a fine degree. After
trajectory determination the CubeSats will perform small course corrections to lineup the pre-
encounter trajectory to attain the flyby characteristics needed and post-encounter trajectory with
the optimum heliocentric trajectory.

The in/out-bound flyby phase attitude will be such that one set of X-band patch antennas
(located on the +z and -z face) are directed toward the available DSN dish when the signal is
being broadcast and the deployed solar panels are pointed in the general direction of the sun.
Saturation of reaction wheels is a major issue during this phase because propulsive maneuvers
are not allowed throughout the flyby. During this phase the trajectory profile will be gathered at
regular intervals by one system. [details needed>>time frame, slew rate profile, DSN dish]
The closest approach phase attitude will be such that the passive SLR reflector (-z face) is
pointed toward the appropriate SLR facility. GPS signals will be received and stored until they
can be broadcast. X-band radio signals may be broadcast to ESA networks to gather additional
velocity information (Doppler). During this phase the trajectory profile will be continuously
measured by multiple systems. [details needed>>time frame, slew rate profile, ESA dish, SLR
facility,]
2.0 Driving Statements and Requirements
This section details the missions scope statements and primary requirements. The
rationale behind each driving statement is included. The result of this section should be a
detailed description of both the limitations of the FLARE mission and the guidelines which will
spur system development.
2.1 Scope
Below is a step by step outline of the scope of FLARE. The need statement should be
considered in reference to our selling statements from section 1.1. The system constraints should
be referenced to mission constraints from section 1.3. The scope it meant to guide/constrain the
project in order to maintain clear and achievable goals and objectives.
2.1.1 Need
Since, so far, the hyperbolic flyby anomaly has defied a full accounting, the question of
whether the anomaly is a real physical phenomenon remains. It is difficult to prove what forces
may be causing the anomaly without a hypothesis to test. Since all previous hypotheses have
been ruled out by accounting for the scale strength of potential perturbations, no likely
hypothesis remains to test. The remaining options are to attempt to prove that the anomaly is a
real physical phenomenon, and then to further characterize the anomaly. Since the
phenomenological formula describing the anomaly’s effects is based on singular data points that
have not been repeated, the first option is both easier to achieve, and would assist the latter
option. Therefore, the need established in this proposal is the following:
To evaluate whether the hyperbolic flyby anomaly is a consistent, repeatable
phenomenon, or an otherwise unaccounted for data artifact.

2.1.2 Goal
To prove the hyperbolic flyby is a real phenomenon, the first step is to show that it is
repeatable. Repeatability requires not only that multiple flybys show anomalies, but that two
flybys of similar or identical characteristics show the same anomalous change in orbital energy.
The phenomenological formula states that the ratio of the change in orbital energy to the absolute
orbital energy is proportional only to the difference in the cosines of the declination of the
incoming and outgoing hyperbolic asymptotes. The change in orbital energy is equivalent to the
change in velocity at the Earth’s sphere of influence, V∞. To show that the anomaly is
repeatable, multiple flybys must be performed with nearly the same declination change. To
further characterize the anomaly and confirm the proportionality constant of the
phenomenological formula, multiple flybys of varying changes in declination must also be
performed and monitored. Therefore, the goals of the proposal are twofold:
To collect a quantity of at least 4 data points during hyperbolic flybys, showing
repeatability of the anomaly, and characterizing its effects.
2.1.3 Objectives
More specifically, the mission intends to supply repeatable data similar to flyby of the
NEAR satellite. The simplest way to accomplish this is to fly two identical spacecraft in very
nearly the same trajectory, with one following the other relatively closely. In addition, the
anomaly can be characterized by repeated flybys with these two spacecraft, varying the
parameters of the joint flyby somewhat between each pass. In order for the flybys to be useful in
analyzing the flyby anomaly, precision tracking data must be acquired for each satellite. In
keeping with the goals, position, velocity, and acceleration data must be collected in a manner
that will allow validation of the previous hyperbolic flyby observations. The mission objectives
are states as:
Collect position, velocity, and acceleration data over the course of at least 4 hyperbolic
flybys from two spacecraft comparable or superior to the data from the NEAR spacecraft Earth
flyby.
2.1.4 Mission
To perform the flybys needed to collect data, two satellites will depart from Earth with a
hyperbolic excess velocity such that the heliocentric orbit will share parameters with the Earth,
except that the inclination shall be increased. Such will encounter the Earth along the
conjunction of their orbital planes, and thus twice a year. The satellites will be tracked and their
kinematic data collected and analyzed to confirm that the anomaly is or is not repeatable and
conforms or does not conform to the current phenomenological formula.
Confirmation and characterization of the flyby anomaly has many potential benefits.
Among them are improvements to the trajectory modelling of flybys, which may increase
available mission possibilities by allowing mission planners to better anticipate the position of

smaller bodies in the solar system. The mission also has the potential to expose the need for
fundamental changes in human understanding of physics.
2.1.5 System Constraints
This subsection is comprised of bulleted summaries and a more detailed description of
broad level constraints. These constraints have procedural, timeline and managerial impacts
primarily. Other constraints are instilled by the mission and system requirements, those reflect
constraints more onto the physical system.
•Projected satellite lifetime (2-4 years) and mission assurance.
–Radiation toll and propulsion capacity.
–250-300 m/s DV corrections capable with 4u worth of hydrazine propulsion.
–Medium to High TRL and rad hardened subsystem components only.
*Redundant systems are a possible substitute for rad hardened systems.
This mission will be limited by the lifetime of the space bound system’s components.
Trajectory correction maneuvers will be necessary to provide trajectory correction maneuvers in
order to maintain recurrent flybys of Earth. Our baseline trajectory is optimistic in terms of
magnitude of the necessary maneuvers and what our system is prepared for. Barring a propellant
leak our system will have far more propulsion capability than demanded by our baseline. One
major assumption in this regard is that our launch system(LV and SHERPA) will provide
sufficient DV to escape Earth’s influence and excess velocity of ~1 km/s.
A more severe limiting factor in this case is the radiation toll on our space bound system.
Although the baseline trajectory provides for rapid transit of the Earth’s magnetosphere (and the
Van Allen Belt’s intense radiation), the satellites will be exposed to continuous solar radiation at
approximately the intensity at 1 AU distance from the Sun. To provide mission assurance either
rad hardened components or redundant systems will be required. Rad hardened systems procure
a significant increase in cost, while redundant systems result in extra volume being taken and
mass increasing.
A final means by which to increase the system’s lifetime and mission assurance is to use
high TRL components. This will decrease testing costs and serve to provide a high confidence
of survivability and capability. Considering cubesats (taking similar precautions and exposed to
similar radiation conditions) in general, the system can be expected to last between 2 and 5 years
barring an unlikely circumstance that wasn’t in consideration (e.g. solar flare, debris impact,
etc.).
•Secondary payload considerations.
–Satellites must be compatible with a Planetary Systems CSD.
–Satellite mass: 10-15 kg. Max satellite volume: 6u.

Figure 3: CSD dispenser typical deployment setup for several 6u scenarios, courtesy of
Planetary Systems Corporation [4], discount lower-right graphic.
The deployment system will be a 6u Planetary Systems capsulized satellite dispenser
(CSD, depicted in Figure 3). The particular CSD to be used is denoted as the 2002367B payload
spec for 6u cubesats. To be compatible with the CSD the cubesat will need two tabs tab running
the length of the cubesat to interface with the deployment mechanism, the -Z axis contacts the
ejector plate (400N force during launch due to vibration) and optionally an electronic interface
on the +Z (prefered) or +X/+Y face for the Separation Electrical Connector (safe/arm plug) [27].
By limiting the size and mass of our CubeSats, the launch associated costs should be minimized.
Although we have “additional launch system needs” (e.g. SSPS), potentially our s/c could be a
secondary payload on that as well, and thus the cost would be shared between parties.
•SHERPA must be compatible with the launch vehicle

Figure 4: SHERPA mounted on a primary payload of a LV [25].
The secondary payload considerations serves to maintain the compatibility of the cubesat
deployment system (CSD) to the SSPS, so the only remaining concern is that (SHERPA) the
launch assist system, is compatible with the LV. SHERPA (also referred to as SSPS) has been
designed to the specifications of medium and intermediate class LVs (as depicted in Figure 4)
such as Falcon 9, Antares and Evolved Expendable Launch Vehicle (EELV) [25]. The particular
SSPS that accommodates the baseline trajectory is the SHERPA 2200, which can produce ~2200
m/s of DV with a 300 kg payload and ~2600 m/s DV with a 30 kg payload [3].
2.1.6 Assumptions
FLARE makes several assumptions that are acceptable and relatively commonplace
assumptions when developing a project. For example, it is assumed that as a secondary payload
our baseline trajectory parking orbit can be acquired. The SSPS is assumed to be in the launch
associated costs category wrt FLARE’s budget and potentially involved in ride sharing to
minimize cost. Although ongoing trade studies point to FLARE being able to achieve the
driving requirement of velocity accuracy, it is assumed that the instrumentation necessary will fit
on a duly (wrt the subsystem requirements, e.g. with the components listed on the PBS) capable
6u cubesat. Although it has been considered as a possible ConOps by several resources (JPL and
others), a highly eccentric orbit was eventually assumed to not have a measurable (wrt cubesat
capabilities) anomaly associated with its closest approach. Some resources (references provided
by JPL contacts) have laid claim to solving the anomaly in one form(high order gravity terms
stacking) or another(anisotropy of the speed of light), FLARE is operating under the assumption
that more data on the anomaly is beneficial to the scientific community.

2.1.7 Authority and Responsibility
The principal investigator for this mission proposal provided the suggestion for the
mission to NASA’s Jet Propulsion Laboratory. As a result, it is NASA JPL that possesses
authority over the mission should it be selected for further development. In such case, JPL
would assume authority over the final development, fabrication, procurement, integration, and
maintenance of the spacecraft. They would also be responsible for the safety of the mission, as
well as flying and ensuring the collection of necessary tracking data.
The University of Texas at Austin student team consisting of Jeffrey Alfaro, Kyle
Chaffin, Anthony Huet, Amritpreet Kang, and Graeme Ramsey, currently known as Team
FLARE, is responsible for the preliminary systems engineering, design, concept of operation,
trade studies, and this proposal.
2.2 Primary Requirements
This section details top level requirements accompanied by a brief rationale. These
requirements are intended to drive the acquisition of data to prove the existence of a velocity
anomaly during flybys (gathering data prevalent to characterizing the anomaly is a bonus). It has
been divided into two subsections, one related to the broader mission and the other focused on
the actual system and its implementation. See Appendix III for lower level requirements.
2.2.1 MissionRequirements
[A] The system shall be capable of measuring a change in orbital energy to the level of
precision of tenths of a millimeter per second changes in hyperbolic excess velocity.
This requirement is paramount to the success of FLARE. Viable data return on the
anomalous velocity change is the directive of this project. Past missions that were able to
accurately measure this anomalous velocity change are referred to as heritage missions These
missions were large scale (microsats and greater in size) whereas FLARE is a secondary payload
with severe size and performance limitations which will make our required measurement
accuracy more difficult to achieve than the heritage missions. This difficulty is due to
diminished volume allowing less capabilities in regards to its components [from power available
to pointing accuracy, this is particularly noted in regards to our perspective GPS device, the most
accurate of which are too large for a 6u cubesat].
[B] This project shall provide at least 4 data points associated with the flyby phenomenon in
its projected lifetime.
In order to make any real conjectures unto the anomaly’s source or further refine the
phenomenological formula a large enough set of data is essential. Considering all known
heritage missions, only 7 data points currently exist. By accruing 4 more data points the
resolution of the data and resulting analysis is almost doubled. 4 data points are achievable in
both of our primary and secondary ConOps.

[C] The system shall be capable of tracking the velocity/position of each satellite throughout
the flyby to tenth-of-millimeter per second/centimeter accuracy.
This requirement serves to further characterize the anomaly. During closest approach
during a flyby there can be a 4 hour gap in trajectory monitoring due to the fact that the DSN
dishes cannot slew fast (need value, what other means for closest approach) enough to track
during that high relative speed segment. GPS monitoring will be able to fill in the gaps of
position and velocity data, though most likely less accurately than required to satisfy
identification of the anomaly. If the accuracy is sufficient to identify the anomaly around closest
approach, it will greatly serve to further our knowledge of the characteristics of the anomaly.
Predominantly, it appears that the anomaly’s source takes place near closest approach, so any
further resolution on the intricacies of the formation of this anomaly will serve to facilitate our
conjectures (phenomenological formula and anomaly source).
[D] The mission design shall perform velocity data collection on at least two “paired” flybys
(with very nearly the same change in orbital energy) at a level of precision of 0.1 mm/s changes
in hyperbolic excess velocity.
This requirement reiterates the most dominate requirement of data precision and refines it
to our ConOps. We intend to use tandem, paired flyby formations to demonstrate repeatability.
Repeatability or deviation from repeatable will further serve to characterize the anomaly. To
identify the anomaly, 0.1mm/s resolution in the measurement of the inbound and outbound
hyperbolic excess velocity is required because the anomaly is expected to be on the order of
several mm/s.
2.2.2 System 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.
This further details primary mission requirement [C], the justification is the same. This is
simply how we intend to implement that requirement. Other viable options for closest approach
coverage include Satellite Laser Ranging (SLR), and Radio Doppler analysis using networks
other than DSN. Position profile data can be differentiated to gather additional complementary
velocity profile data. Multi-platform and cross-platform (e.g. differentiating position data to
velocity while also gathering velocity measurements using one platform) velocity tracking, that
is to say “gathering multiple independent velocity profiles”, is not a listed requirement, but
would increase mission assurance and data confidence if implemented and should be considered.
{B} Confirmation of an anomalous DV shall be achieved via Doppler effects from X/S-band
radio broadcasting during the flyby phases.
This serves to satisfy our need for velocity measurements over most of each flyby
trajectory, thereby identifying if there was a measurable anomaly. DSN will be responsible for

gathering the velocity profile other than closest approach (where the slew rate of DSN dishes
prevents coverage).
{C} The error of Doppler velocity measurements shall be on the order of 0.1 mm/s.
This satisfies primary mission requirements [A] and [D]. This order of accuracy has been
achieved in our heritage missions using similar bandwidths(X-band) and technologies(which
have been/are being scaled down to cubesat specifications).
{D} The satellites shall be constrained to a standard 3u/6u CubeSat format.
By minimizing the size of our satellite, the budget of the overall project is reduced. This
size restriction also serves to provide a baseline for capabilities and constraints regarding
implementation and performance.
{E} The satellites shall perform flybys with sufficient hyperbolic excess velocity and change
in declination to produce a predicted anomaly of at least ±3 mm/s.
This assigned minimum of the expected anomaly for each flyby assists in trajectory
design. It is an appropriate value inline with what flyby characteristics the baseline trajectory
predicts. It also serves as a complement to the proposed velocity data accuracy such that a
healthy margin is maintained to assure a confident anomaly identification. Our baseline
trajectory provides a predicted anomaly of over 5 mm/s for each flyby.
{F} The altitude of periapse upon each flyby shall be between 500 and 2000 km.
The phenomenological formula fits flybys with periapse between the above altitudes.
This requirement is intended to assure the predicted anomaly is accurate and by that standard
maintain confidence that the anomaly would be measurable on that trajectory if it does exist.
The lower bound of 500 km will keep the satellite from experiencing noticeable atmospheric
drag. Whereas the upper bound simply marks where the phenomenological formula starts
experiencing higher error wrt the heritage mission data. The baseline trajectory will aim for a
distinct periapse altitude between 500 and 2000 km for each flyby, the particular altitude itself is
not important and was a variable in optimizing the trajectory.
2.2.3 Requirements Traceability Matrix
The primary mission and systems

Table 3: Primary Requirements Traceability Matrix, including mission requirements not
explicitly listed in section 2.2.1, after the label [extra].
3.0 System DesignDevelopment
This section will describe the steps taken prior to developing our Midterm Design. It
serves as a description of the first iteration of mission and system development via research and
trade studies. Addressed below are the most important factors in the early goings of FLARE’s
development. These factors include: ConOps and scope refinement to drive the mission,
creating a baseline trajectory to prove feasibility, producing a baseline PBS to spur further
component research, researching the proposed data acquisition systems (GPS, radio Doppler),
and accumulate significant design heritage. These and other trade studies allowed the
recognition of critical parameters to drive the remainder of the project (summarized at the end of
this section).
3.1 DesignAlternatives Development
In our preliminary brainstorming and researching into the flyby anomaly we produced 3
different ConOps scenarios. These ConOps scenarios had varying characteristics as to what
quality and quantity of data they returned, along with cost and timeframes associated with the
mission. ConOps B served as our baseline ConOps scenario after preliminary evaluation.
This scenario involves multiple cubesats (>2) on highly eccentric elliptical orbits around
Earth. Each satellite would follow a trajectory at a different declination. It was assumed that the
anomaly might be observable in highly elliptic orbits. The satellites would perform these orbits
to see if the anomaly was notable in captured orbits. After a large number of captured orbits, the
satellites would perform a DV maneuver to then be set upon a hyperbolic trajectory and attempt
to measure the anomaly. This option produced an unsure amount of data (due to unknown
quality), in a very short time frame for low cost.
This idea was ruled out for several reasons. First, according to the phenomenological
formula and available data, the magnitude of the anomaly is scaled with velocity and thus the
sensed anomaly would be miniscule to non-existent for captured orbits. Second, the
phenomenological formula and available data point out that a sufficient change in declination is
required on inbound and outbound legs, this translates to a plane change for captured orbits
which is difficult to achieve. Finally, this idea lacks merit due to the fact that a DV near periapse
would disallow a certain measurement of the anomaly.

The second scenario involves a single flyby event using a “mothership” and many (~6)
3u cubesats. The mothership with cubesats docked would be perform an EVE boosting
trajectory. Upon approach of Earth after Venus, the cubesats would be deployed and perform
paired flybys at varying parameters to demonstrate repeatability for multiple circumstances.
These cubesats would essentially be sensors (GPS and X-Band Doppler for data) with the ability
to perform small DV maneuvers and ADCS pointing while within ~0.1 AU of Earth approach.
This option produced a large amount of data of great quality, in a medium time frame for
medium cost.
With the boost from Venus our satellites would have sufficient excess velocity with
respect to (wrt) Earth such that the predicted anomaly would be on the order of 10 mm/s. This
would decrease the needed sensitivity of the systems instrumentation or alternatively increase the
resolution of the anomaly, aiding to refine the phenomenological formula. Seven (including
“mothership”) data points would be provided in a relatively short time period. This scenario has
been molded into the primary ConOps (described in section 1.5). The most notable changes
being a shift to multiple Earth flybys using two 6u CubeSats.
The third ConOps scenario is a recurrent flyby event using one relatively capable
microsat. This microsat would perform a variety of heliocentric maneuvers to produce multiple
Earth flybys, starting with an EVE maneuver to boost energy. This microsat would be much
more capable than the cubesats considered in all other ConOps. It would incorporate multiple
means of accurate velocity profile acquisition, and possibly other instrumentation in an attempt
to characterize the anomaly or rule out some proposed causes. This option produced a low rate
of data return of extremely high quality and high cost and was ruled out accordingly.
This idea maintains merit if piggybacking on a mission is possible. Meaning, if a current
mission had planned a flyby of Earth which would follow a trajectory providing a measurable
(measurable given the satellite’s instrumentation and Earth ground support) predicted anomaly,
the velocity profile could be applied to the analysis of the anomaly. One such mission was
JUNO (see section 1.3.5) from which a velocity profile including closest approach was produced
after it performed an Earth flyby in 2013.
3.2 System and Subsystems Allocation
After settling on a ConOps which would require either a 3u or 6u cubesat format, a
preliminary Product Breakdown Structure (PBS) was created to guide the investigation into
component selection. Throughout the design process the preliminary PBS evolved into a mature
form depicted below in Figure 5. One early design consideration was the propulsion system.
Hydrazine was the first choice for cubesat propulsion system due to its high DV capabilities.
Secondary payload considerations due to the toxicity/volatility of hydrazine render cold gas or
electric propulsion as substitutes (with less DV capability). Hydrazine was decided upon as the

best system after our JPL correspondent advised that it was an acceptable risk and not
uncommon in recent launches. The largest point of contention was and continues to be the
selection of components which are the source of data acquisition in regards to the anomaly. The
first design choice included dual frequency X/S-Band radio and a dual frequency L1/L2 GPS
receiver. The more mature design choices use a JPL developed X-Band transponder and also has
GPS outlined in red to signify it might be replaced with SLR (via a passive reflector). The items
outlined/highlighted in red may either be replaced with a comparable system (propulsion) or
dropped entirely (TPS).
Figure 5: FLARE Primary ConOps PBS, orange = primary to mission anomaly data, yellow =
datasource, red = in contention.
3.3 System DesignHeritage
This section describes the approach used and heritage acquired to design our system.
Dominant heritage is depicted in figures, primarily data acquisition systems and “semi-deep
space” (outside of Earth’s orbit) cubesat system architecture.
3.3.1 INSPIRE Cubesat
JPL’s Courtney Duncan produced several presentations in regard to Iris (X-band Comms
system) which have proved invaluable [33-35]. The INSPIRE cubesat (depicted in Figure 6) was
the first to leave Earth orbit, its system will be very similar to the systems needed by FLARE.

Not only are components listed and depicted, a brief overview is provided showing the basic
characteristics and capabilities of the cubesat.
Figure 6: INSPIRE cubesat provided for subsystem design heritage [33].
3.3.2 X/X-band LMRST
This JPL developed X-band radio system demonstrates the components that will go into
FLARE’s Comms subsystem. Another Courtney Duncan (of JPL) presentation regarding Iris
provided this example of cutting edge of CubeSat Comms. The Low Mass Radio Science
Transponder (LMRST) depicted below in figure 5 is a 2014 model, 1u in size, ~1 kg in weight,
demanding 8 W when active, and capable of achieving 1 m accuracy ranging. The goals listed
for the immediate future in regards to LMRST capability are 0.5u size, 3 W power when active,
with an approximate cost of $100,000 for a unit. [34]

Figure 7: JPL developed Low Mass Radio Science Transponder with X/Ka options [34].
3.3.3 Iris X-band Transponder
A second potential X-band transponder configuration is depicted below in Figure 7. To
reiterate this is the most important system for FLARE as it is the primary source for
identification of the anomaly’s presence. The Iris (not an acronym) transponder depicted below
is 0.4u in volume, 400g in mass, and requires 10 W of power when active.

Figure 8: Iris X-Band transponder system, courtesy of JPL [33].
3.3.4 GPS/GNSS Receivers
When examining GPS receivers that would potentially provide post-processed velocity
accuracies of millimeters per second, the “BlackJack” GPS Receiver (Figure 9) developed by
JPL demonstrated the capabilities that a space based GPS receiver could achieve on missions
such as GRACE, JASON-1, and CHAMP. Unfortunately, due to the mass and volume
constraints of the FLARE mission, the BlackJack GPS Receiver was not a viable option for this
spacecraft.
Figure 9: BlackJack GPS Receiver, courtesy of JPL[38].

Figure 10: Radio Aurora eXplorer (RAX) CubeSat [43].
[discussion of Figure 10 needed]
Additional receivers that were considered include the FOTON receiver proposed by The
University of Texas at Austin and various receivers manufactured by NovAtel. Single frequency,
L1 GPS receivers were considered and then ruled out due to their low accuracy.
3.4 Trade Study Summary and Results
After defining the baseline system design, several trade studies became necessary to
advance the project further. The most important trade studies wrt the mission goals and
objectives are related to the data acquisition systems and trajectory design. Other important
trade studies with broad design ramifications include a launch vehicle and parking orbit
characteristic trade study, a propulsion system trade study and an evaluative trade study between
the two ConOps in contention for primary. This section will describe those evaluations and the
thought processes associated with it.
3.4.1 Data Acquisition Systems
A large variety of resources were accumulated in reference to radio Doppler analysis and
Comms systems in cubesats. Most helpful and abundant of these resources were discussions by
JPL’s Courtney Duncan. Her papers and presentations [33-35] provided great insight into the
current state of the art in regards to cubesat Comms and their use for GN&C. Figure 10 below

helped rule out Ka-Band as a candidate component, seeing as X-Band patch antenna data rates
were sufficiently large at the ranges expected for our data gathering (<0.0062 AU) and ranges
expected for our trajectory correction commands (<0.008 AU).
Figure 11: Radio band comparison for cubesats, courtesy of NASA JPL [19].
Most of the heritage missions observed the anomaly by use of X-Band radio Doppler (all
by some form of radio Doppler) analysis
Several resources were accumulated in reference to GPS accuracies [15-18, Amrit can
you add relevant GPS resources here?], and in particular velocity accuracy in regards to post-
processing. Listed in Table 3 below are steady-state navigation errors after 23.5 hours of
trajectory processing, “i.e. the filter has converged to a minimum error with consistent covariant
estimate” [21]. The values in Table 3 apply to Goddard Space Flight Center’s PiVoT GPS
receiver with weaker signals from 28 to 25 dB-Hz [21]. It is worth noting that this report is from
2001 and advancements in the field of CubeSats are bound to have increased CubeSat GPS
capabilities.
Seeing as FLARE has no need to calculate real-time trajectory profiles, the steady-state
values are assumed to be representative of the level of accuracy achievable in post-processing.
The only part of the flyby phase GPS will need to cover is the section where the satellites are
moving at an angular rate beyond the slew rate capabilities of DSN. [slew rate discussion
needed, and altitude of GPS or closest approach coverage]

Table 3: steady-state GPS navigation errors [21], for analysis of expected accuracies. Two
perigee passes were necessary to achieve this level of steady-state accuracy.
The GPS equipment [21,38,Amrit can you list all your relevant GPS resources here?]
used is an ultra low power receiver designed specifically for small satellites. Due to the nature of
the mission, it is imperative that the GPS unit be reliable and provide accurate data, which this
unit is well tasked for. It will begin operating within 5 minutes of activation, and has no altitude
or velocity limitations. A significant feature of this unit is the ionizing radiation shield. Since the
spacecraft will be travelling outside of the Earth's protective magnetic field it is necessary to
have radiation protection, more so than for typical LEO missions. NASA and ESA preferred
component vendors are used as suppliers and finally it is assembled in an ESA certified 100.0
clean room. Overall this GPS unit has many qualities that makes it an excellent choice for this
mission.
3.4.2 Launch Vehicle
Determining if the Russian launch vehicle, Rokot, was a viable candidate for our system
given its circumstance of being a secondary payload was a preliminary investigation coupled
with the baseline trajectory needs. Traditionally Rokot delivers its payload to 500-1000 km
altitude and in the process varying its flight path angle such that it will circularize the orbit. A
simple way to approximate if any given circular orbit was a viable scenario given the means of
Sherpa 2200 as the launch assist vehicle is depicted in Figure 11. This figure allows for
visualizing the velocity maneuver (DV) necessary (modeled as an impulsive burn) to escape
(with no excess velocity) Earths influence from a circular orbit, and the maximum excess
velocity providable by a Sherpa 2200 (under minimum and maximum load) again assuming an
impulsive burn from a circular orbit.
From first glance it is apparent that Rokot under standard launch procedures is not a
viable solution even under minimum payload conditions (excess velocity of ~ -450 m/s, e.g. still

in a captured orbit). The option remains available to given a Rokot launch which doesn’t
circularize the orbit would allow the DV maneuver to be performed at periapsis of an elliptic
orbit (a much more efficient procedure). A circular orbit our only available parking orbit, in
order to achieve an excess velocity of 0.5 km/s an altitude of 9000 km would be necessary. This
should be enough evidence that FLARE cannot launch into a circular LEO, and launching into a
circular orbit at all seems like a waste of SSPS fuel.
The result of this trade study along with the trajectory trade study shows that as opposed
to Rokot, an intermediate class launch vehicle like Falcon 9 is a viable option. Essentially the
Trajectory trade study demands a highly eccentric (>0.7) and inclined (~60 deg) parking orbit
with a semimajor axis near 25,000 km which reinforces an intermediate class launch vehicle as
the best option. Listed on Space Flight Services are several 2018 launches destined for highly
eccentric and inclined trajectories. In particular several Russian launches were destined for HEO
at ~60 deg inclination, these could fulfill our launch vehicle requirements.
Figure 12: MATLAB coded Rokot LV analysis, in conjunction with SHERPA 2200, circular
orbits, impulse DV.
3.4.3 Trajectory
[a discussion of our how we came to decide on our trajectory needs to go here, also the
velocity triangle of our departure orbit that guided us initially, maybe a depiction of the satellite
departing from earth which is tilted in solstice w the satellite traveling to the +z direciton wrt
orbital plane]
A preliminary trajectory for the primary ConOps was found using TRACT (described in
baseline section) that meets the mission constraints. Trade studies to optimize the spacecraft
trajectories are TBR. The intent is to determine if further flybys can be achieved without large

DVs. Furthermore, the final leg of the spacecraft’s trip should be evaluated to determine
disposal options.
The secondary ConOps also requires a full trajectory workup along the lines of that
performed for the primary ConOps. Such a study would allow a more comprehensive
comparison of the two options.
3.4.4 Primary (A) vs. Secondary (B) ConOps
Development of the secondary ConOps baselines and thus an empirical evaluation of the
merits of each mission approach and selection is TBR. This section will detail our preliminary
assumptions that lead to the primary ConOps selection.
Table 4: Design selection criteria and weight, for ConOps (primarily) and system evaluation.
The criteria listed above in Table 4 serves to enumerate the importance of each criterion.
Anomaly magnitude refers to the expected anomaly via the phenomenological formula. Budget
refers to the entire mission costs, from mission development to launch and maintenance costs.
Data Quantity refers to the quantity of velocity profiles (or anomaly data points) in the expected
mission lifetime. Turnover Ratio refers to the rate of data return, it is represented as expected
data points divided by mission time. Mission Assurance refers to the level of confidence that the
mission requirements will be satisfied.
The evaluation of each ConOps based on empirical means for what became the primary
ConOps and non-empirical means for what we retained as the secondary ConOps is summarized
below in table 5. The rationale for this decision process is listed below. An empirical study of
the secondary ConOps and reevaluation of the design selection is TBR.
•Maximized anomaly magnitude (>3mm/s)
The anomaly magnitudes must be sufficiently (at least an order of magnitude) greater
than the error associated with the system instrumentation. This is the most important factor as it
defines the quality of data that FLARE must retrieve. The phenomenological formula is the
basis for quantifying the anomaly, however it is only an estimate thus the anomaly could be
smaller than expected. A velocity measurement error of 0.5 mm/s and an expected anomaly of 3
mm/s (minimum) will serve to supply a marginally sufficient situation. Either increasing the
expected anomaly or decreasing the error of the velocity data acquisition system serves to better
satisfy this parameter.
•Minimized Budget (<$5mil)

Budgetary constraints are an important constraint. Our budget limit is currently set at
$5mil excluding launch associated costs. This parameter refers to minimizing both our expected
budget and launch associated costs. [costing overview needs to go here]
•Significant Data Quantity (~4 data points)
This parameter represents the second mission requirement [B], which is marginally
satisfied by a system that provides 4 data points in its projected lifetime. This factor is slightly
less important than most of the others listed here. The primary ConOps provides for 6 data
points (velocity profiles that have an expected anomaly). The secondary ConOps allows for 4
data points.
•Rate of Data Return (~2 data points per year)
This is the least important factor to the success of FLARE (called Turnover ratio in
tables). Although less time means less management costs, the rate of data return is not
paramount to the overall mission goals and objectives. 2 flyby anomaly data points per year (or
0.1666 data points per month) describes as marginally sufficient condition. The primary ConOps
gathers 6 data points in 2 years, this gives it a ratio of 0.25 data points per month. Whereas the
secondary ConOps gathers 4 data points in less than 2 months, giving it an approximate ratio of 2
data points per month.
•Mission Feasibility (~mid/high TRL and low risk)
Just as important as budgetary consideration, a system must maintain feasibility through
mission assurance. Without readily available technologies and tested systems mission assurance
diminishes. For purposes of comparison, mid to high TRL subsystems and low risk to mission
assurance was defined as marginally sufficient. [risk and system integrity details needed]
Table 5: Tentative design selection results, A = Primary ConOps, B = Secondary ConOps.
•Repeat tandem flybys of Earth
The Primary ConOps won this tentative evaluation as depicted in Table 5 above. As a
result the particular system associated with the primary ConOps will be the focus of FLARE’s
efforts. Once further evaluation, in particular a baseline trajectory, is provided for the secondary
ConOps, the system capabilities will be defined and the selection criteria can be weighted by
derived values instead of assumed values.
•Choice based on precursory characteristics
The Primary ConOps use of multiple flybys serves to boost the anomaly magnitude
consecutively. This factor along with the fact our baseline trajectory has provided evidence that
our expected anomaly will go from over 50 to over 70 times greater (wrt the first and second
baseline flybys) than the projected system velocity accuracy. [more evaluative details needed]

In tentatively evaluating the Secondary ConOps, the main factors that could for certain be
in its favor are data quantity and rate of data return (turnover ratio). The expected duration of
this mission would be mere months, and with potentially 10 or more 3u cubesats being deployed,
this approach definitely has merit. The cubesat system wouldn’t need to be as capable (less
propulsion, no rad hardening, etc.) and thus each cubesat would cost less saving on cost.
However without a baseline trajectory the launch associated costs which dwarf other costs is to
be resolved. [more evaluative details needed]
•Verification/analysis of assumed characteristics TBR
3.4.5 Propulsion
Several potential propulsion systems were considered for use on the spacecraft.
Ultimately monopropellant hydrazine motors were decided on due to their high TRL level and
ease of integration into the spacecraft. Hydrazine also provides high thrust, which simplifies the
trajectory calculations by allowing the mission designer to consider space burns to be relatively
impulsive.
Other contenders were electric propulsion, bipropellant engines, and solar sails. These
were considered with the goal of reducing propellant mass. Additionally, alternative propulsion
methods were considered due to the need for ride-sharing. If the spacecraft are to be a secondary
payload of a launch, the primary payload operator may object to potential contamination from
hydrazine propellant and outgassing.
Electric propulsion systems such as ion engines have high specific impulse, but
unfortunately lack the thrust levels desired for this mission. Since the thrust maneuvers must be
executed in a relatively short amount of time, current electric propulsion systems would not
provide sufficient thrust to carry out the mission. In addition, many current electric propulsion
systems lack the TRL to be used in this mission and would add too much risk to be deemed
worthwhile. [I thinks Hydrazine and Electric Propulsion are the two best options for primary
ConOps, for the secondary ConOps Cold Gas or a small Hydrazine motor would be best.]
-not a criticism, just my thoughts on the project, perhaps should be mentioned
Bipropellant engines offer high thrust and moderate specific impulse levels. However,
bipropellant engines on this size of cubesat have not been fully developed and integrating a new
propellant system is not worth the added risk.
Another option was solar sails. However, these have the lowest TRL of any of the options
available. These also have the similar problem as electric propulsion in that they provide very
low levels of thrust. In addition, since the flyby must be unpowered in order for the anomaly to
be measurable, the solar sail would have to be detached sometime prior to the flyby event
(Earth’s SOI), further complicating the mission.
Monopropellant thrusters have a long heritage in spacecraft applications. They are also a
relatively simple system that requires only one propellant. While it is the least efficient method
considered, it still provides ample thrust for the spacecraft maneuvers to be completed in a timely

manner. Overall these factors made monopropellant thruster stand out as a clear choice for the
propulsion system.
3.5 Critical Parameters
•Tracking ability during the non-closest-approach phase of each flyby
The FLARE mission’s success depends upon tracking cubesats during flybys of Earth. If
the cubesats are not trackable, the mission will fail. The goal at this phase of the trajectory is to
find the inbound and outbound excess velocities and gather enough trajectory information to
build an accurate trajectory profile. Pointing requirements are designed to accommodate ground
stations such that the X-band radio signals from the spacecraft produce the most accurate
velocity profile. JPL midterm feedback revealed the fact that a tumbling satellite’s velocity data
can be just as accurate or more, in post processing. This fact deserves further consideration.
As section 1.5.4 details, during the flyby the satellite will maintain an attitude to point at
a DSN dish until the closest approach phase. This entails that the attitude control system must
avoid saturation over the approach and departure legs of each flyby. One consideration is to use
torque rods to desaturate the reaction wheels during the closest approach phase to prepare for the
outbound leg.
Lastly, NEN and DSN availability is critical to the tracking ability of the spacecraft.
[details of DSN usage patterns, exactly how much of the (non-closest approach)velocity profile
is adequate for our mission, and which dishes our baseline trajectory can use]
•Tracking ability during the closest-approach phase of each flyby [NEW]
The section of the trajectory around periapse of the flyby where the DSN slew rate
disallows monitoring of the CubeSats is defined as the closest-approach phase. This is the area
where 6 of the 7 heritage missions lack coverage. The anomaly seems to take place near
pariapse, according to trajectory propagating models (JPL) the inbound and outbound legs of
those 6 heritage missions are discontinuous at periapse, represented as an anomalous change in
velocity. In reality the effect must be gradual, regardless, the closest approach phase is the most
important section of the trajectory in regards to data that could be used to characterize the
anomaly, not only identify it.
A variety of instrumentation has been considered for closest approach coverage.
Multiple means of coverage would serve to strengthen data confidence and is a consideration.
GPS was the initial consideration for primary system during this phase. X-band Radio Doppler
coverage during closest approach was demonstrated during the JUNO flyby with collaboration
between JPL and the European Space Agency (ESA). This means would be more accurate than
GPS and wouldn’t require another subsystem thus it is the top contender. Satellite Laser Ranging
(SLR) is the best complementary system for our mission, the only additional component is a
passive reflector. SLR would gather very accurate position data which would be differentiated to
gather a complimentary velocity profile.

•Reevaluate design choice based on an empirical trade study
Between the Primary and Secondary ConOps, referencing section 3.4.4 the primary
ConOps remains the primary choice for the mission thus far. However, an empirical study of the
secondary ConOps has yet to be performed. Once ConOps B has been further developed, and a
baseline trajectory produced, a reevaluation of the design selection is TBR.
•Radiation exposure during heliocentric trajectories
Another consideration that is critical to mission success is the radiation exposure the
spacecraft will be subjected to upon its heliocentric trajectory. The components chosen for the
baseline design have been identified to have a lifetime of two to three years in Earth orbit, as
provided by the manufacturer specifications[Amrit, resources?]. In order to extend the lifetime of
the spacecraft, the components may need to be further radiation hardened or radiation shielding
may need to be added to the spacecraft.
• Vibration during launch
Although most of the components listed in section 4.2.2 have been guaranteed to
withstand certain vibration loads, an analysis of the vibration experienced during launch and
operations has yet to be performed. Upon completion of this analysis, alternate components may
be chosen. [talk about perspective CAD model use for analysis]
•Thermal requirements
The operating temperatures of sample components aboard the spacecraft are given in
Table 6. These thermal constraints limit the operation of the satellite and may warrant the
addition of passive and/or active thermal protection systems. Upon completion of an analysis
consisting of the thermal inputs and outputs to the spacecraft, components such as radiators may
be added to the spacecraft in order to keep components between certain temperature limits.
Additionally, the thermal requirements of each component may dictate the internal layout of the
spacecraft.
Table 6: Thermal requirements, Primary ConOps system design. (needs midterm update)

3.6 Midterm Design Refinement
This section outlines the feedback and design refinements that took place regarding the
midterm presentation at UT and at JPL.
3.6.1 JPL midterm mission design presentation feedback
Below is a brief description of the feedback provided by JPL during our visit to their
facilities and presentation to the below mentioned individuals. To see a detailed accounting of
the valuable feedback provided by JPL associates: Travis Imken, Jackie Green, Randii Wessen,
Bill Frasier, Damon Landau, Jeff Stuart, Macon Vining, John Elliott, Eric Gustafson and Melissa
Vick, reference Appendix IV.
Programmatics
More information regarding why NASA would care about the project was requested.
This was accommodated in section 1.1, denoted selling statements. More details on cost, risk
and schedule were requested along with cost-sharing avenues. These details are TBR.
ConOps
This was one of the primary sources for discussion at JPL. More description of the
primary and secondary ConOps were requested as well as an empirical evaluation of the two
candidate ConOps. An investigation into modeling was prompted, by knowledge of pertinent
perturbation needs. Also Satellite Laser Ranging (SLR) was suggested as a “power passive”
additional sensor.
Baseline/Trade Space/Subsystem
This was the other notably large source of discussion. Many factors from rad-hardened
vs. redundant systems, spin stabilization techniques, range of parking orbit possibilities, to
identification of critical communication errors (mistypes, e.g. transceiver should be transponder).
Also suggested was Surrey Space Systems for GPS and propulsion systems. Other items that
were brought up subsequently were out gassing perturbations affecting the secondary ConOps
and a report on a JUNO flyby of Earth as the most recent heritage mission wrt the anomaly.
3.6.2 Launch Vehicle and Launch Trajectory Details
TBR. Top on the list is fulfillment of a parking orbit which can facilitate the baseline
trajectory. The necessary characteristics of a suitable LV are: delivering its primary payload
(and thus FLARE as a secondary payload) to a highly inclined, highly eccentric trajectory upon

deployment. Also necessary is compatibility with Sherpa 2200, which demands an intermediate
or medium class launch vehicle.
3.6.3 Burn at Earth SOI Calculations
While the TRACT software was capable of optimizing the trajectory down to the level of
patched conics approximations, the expected DV quantity for orbital maneuver corrections is still
needed in order to refine the physical design of the spacecraft. Thus, an estimate of the DV
needed for these ‘known unknowns’ is required. Since interplanetary type CubeSat missions are
not well defined, historical data from other interplanetary missions must be used. JPL’s advice
will be crucial in this estimation, and the figures are TBR.
3.6.4 Subsystem Component Choices
Ka-band radio has been eliminated from contention as a potential Comms subsystem. It
only serves to provide better data rate, and is intended primarily for communication over great
distances. This is beneficial to the overall design as Ka-band requires a substantial amount of
power compared to X/S-band radio which has been chosen as our Comm system’s mode of
information transceiving.
3.6.5 CAD Model for Analysis
A CAD model was made using the battery, flight computer, EPS, power distribution
system, and structure shown in section 4.2.2., in addition to the SGR-05U - Space GPS Receiver
by Surrey Satellite Technology US LLC. and the VHF downlink / UHF uplink Full Duplex
Transceiver by Innovative Solutions In Space. In order to assess the viability of a six unit cubesat
with components similar to those in section 4.2.2., this early CAD model, as seen in Figure 12,
was developed.

Figure 13: Early CAD model for a FLARE cubesat.
In this CAD model, the components discussed above would fit in the two PCB stack
compartments pictured. This modular configuration would allow ample room for the two unit
propulsion system, in addition to the attitude determination and control system.
3.6.6 Final Flyby Maneuver and System Disposal
The final trajectory of the spacecraft, whether into interplanetary space or bound for
collision with another body (Moon impact), or it may burn up upon re-entry of Earth or
otherwise be removed as potential space debris, is not much of an issue and is TBR.
4.0 System Design
See subsection 1.5 for midterm ConOps description as an introduction to this section.
This section will describe the FLARE team’s findings and approach at the end of the project
development cycle.

4.1 Baselines
This section details our preliminary approach and the baselines that the team developed
to guide the project into maturity. It is comprised of component selection baselines.
The Master Equipment List (MEL) serves as a mass budget table for a FLARE
spacecraft. Components were selected for the highest weight to produce a conservative estimate.
This analysis may thus be considered a worst case scenario, with the components shown in Table
1 of Appendix I. This MEL does not include a radiator or any antennae that may be needed for
communication.
The MPS-120XL CubeSat High-Impulse Adaptable propulsion system is a hydrazine
propulsion system that utilizes four thrusters. The BCTXACT is a 3-axis attitude determination
system that utilizes a star tracker, IMU, sun sensor, three reaction wheels, a magnetometer, and
three torque rods in order to determine and control spacecraft attitude. The OEM638 Triple-
Frequency GNSS Receiver serves as a GPS receiver for position determination. The IRIS
Navigation and Telecomm Transponder serves as the radio communication for the FLARE
spacecraft with the Near Earth Network (NEN) and the Deep Space Network (DSN). The ISIS
On Board Computer is a flight computer used to monitor and control all subsystem components.
The FleXible EPS system is an electrical power system that maintains the power systems on
board including the battery, solar panel, and power distribution systems.
4.1.1 Primary ConOps Baseline Trajectory
A baseline trajectory for the primary ConOps was solved for using TRACT, an orbital
trajectory optimization tool developed by Martin Brennan at the University of Texas at Austin.
The trajectory consists of departure from a highly elliptic, and eccentric parking orbit similar to a
Molniya orbit.
A burn at perigee, as can be seen in Table 7, will place the spacecraft into its departure
trajectory, resulting in a V_inf near 3.7 km/s. With the correct launch date to account for the
axial tilt of the Earth, the spacecraft will be placed into a heliocentric trajectory with orbital
parameters that match those of the Earth about the sun, with the exception of a ~7 degree
inclination. Leg 1, as shown in Figure 13, will place the spacecraft on a course to rendezvous
with the Earth in half a year. The flyby at that time, shown in Figure 14, places the spacecraft
onto Leg 2, with an orbital correction maneuver at perigee of 90 mm/s. In order for the flyby to
collect useful data, it must be unpowered, but the orbital maneuver burn in the solution is on the
order of magnitude of error for the patched conic method, so the correction will be within orbital
correction maneuvering contingency. The second leg is slightly more eccentric than the Earth’s
orbit, but with the same total orbital energy. Thus, it will rendezvous for the second flyby after a
period of 1 year. Table 8 gives the orbital parameters of the flybys and their predicted
anomalous energy changes according to the phenomenological formula.

Figure 14: Baseline trajectory, departure (top) and heliocentric phases (bottom) depiction. The
green trajectory is Leg 1, and the cyan is Leg 2.
Table 7: Relevant data for baseline departure and heliocentric trajectories.

Figure 15: Baseline trajectory, first (top) and second (bottom) flyby depiction.

Table 8: Primary ConOps baseline flyby 1 and 2 relevant data.
4.2 DesignChoice
This section will outline FLARE’s midterm system design choices on a system and
subsystem level. The system was chosen to satisfy the mission requirements. Mass, power and
volume considerations are the primary derivative of the design choice and are included after the
overview. Less rationale is provided for the midterm, as opposed to final, report due to the fact
trade studies wrt component choice are ongoing and subject to change.
4.2.1 System and Subsystem Overview
Figure 8 shown below depicts the product breakdown structure of the primary ConOps of
FLARE. The Product Breakdown Structure gives a visual overview of the subsystem allocations
associated with the space bound system. The system is comprised of standard subsystems wrt to
heritage missions with similar trajectories and requirements. Figure 8 is essentially Figure 5
with all the components considered in the following sections (MEL/PEL/EVAL) evaluating the
CubeSat systems mass, power, and volume requirements.
C&DH is comprised of a computer, software and a recorder. The propulsion system is
comprised of a hydrazine motor other choices were ruled out as per the preliminary propulsion
trade study. Power is composed of a solar array, battery and power distribution module. TPS
may only need passive systems, but patch heaters and a radiator are considered. The ADCS
should be comprised of reaction wheels mems gyros and a star sensor, a torque rod could be
useful if saturation is foreseen as an issue during the “quiet” flybys of Earth (severe time
constraint where torque rods can be used=waste of mass). The structural choices are a 6u shell
with interfaces for the CSD (described in section 2.1.5 and depicted in Figure 3), a solar array
deployment system and a SLR reflector. The systems that are most important are highlighted.
The most important systems are the source FLARE’s primary data acquisition, meaning
identification of the velocity profile during flyby phases. The Comms system will be comprised
of an X-band radio transponder, X-band patch antennas (4) along with a UHF radio and
deployable low gain-antennas (2). The Sensor system will consist of a dual (or greater)

frequency GPS receiver, a position and time board, and low-gain antennas previously mentioned.
Further TS will determine the projected capability of this iteration of design choice wrt velocity
data accuracy and the Comms/sensor systems (and possibly SLR).
[Amrit or someone: update PBS to Equipement List PBS]
(PBS is in subsystem folder)
Figure 15: PBS of components considered in the MEL, PEL and EVAL.
Additionally, the cubesat system described above will be accompanied by a deployment
system (CSD), launch assist system (Sherpa 2200), and LV (maybe a falcon 9 or intermediate
class Russian LV) in order to complete the space bound mechanical systems. Additionally there
will be ground based systems such as NEN, DSN, and possibly facilities wrt SLR. Other ground
based “systems” include operation management and on the sideline, the scientific endeavor wrt
analysing the data gathered and investigating not only the phenomenological formula but also the
proposed anomaly sources.
4.2.2 MasterEquipment List (MEL)
For the baseline system design, components were chosen as outlined in the MEL (master
equipment list) seen in Table 8. These components satisfy the requirements outlined in section 2.

FLARE report pre-finalized version

FLARE report pre-finalized version

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