FLyby Anomaly Research Endeavor

FLARE Final ReportGraeme Ramsey, Jeffrey Alfaro, Amritpreet Kang, Kyle Chaffin, and Anthony HuetMay 08, 2015ASE 374L Spacecraft/Mission Design: Dr. FowlerThe University of Texas at AustinIn conjunction with JPL: Travis Imken and Damon LandauSpring 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 Introduction1.1 Heritage

1.1.1 Phenomenological Formula (NEW)1.2 Mission Motivations1.3 Unconfirmed Explanations of the Flyby Anomaly

1.3.1 Dark Matter Encircling the Earth1.3.2 Modifications in Inertia1.3.3 Special Relativity [needs anisotropy of light discussion]1.3.4 Lorentz Accelerations1.3.5 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 Requirements2.1 Scope

2.1.1 Need2.1.2 Goal2.1.3 Objectives2.1.4 Mission2.1.5 System Constraints2.1.6 Assumptions2.1.7 Authority and Responsibility

2.2 Primary Requirements2.2.1 Mission Requirements [slew rate sat & DSN needed]2.2.2 System Requirements2.2.3 Requirements Traceability Matrix (NEW, needs discussion)

3.0 System Design Development3.1 Design Alternatives Development

3.1.1 Preliminary ConOps 13.1.2 Preliminary ConOps 23.1.3 Preliminary ConOps 3

3.2 System and Subsystems Allocation3.3 Design Heritage

3.3.1 INSPIRE CubeSat3.3.2 X/X-band LMRST3.3.3 Iris X-band Transponder3.3.4 GPS/GNSS Receivers [addition figure, discussion needed]3.3.5 Satellite Laser Ranging (NEW)

3.4 Trade Study Summary and Results3.4.1 Data Acquisition Systems [slew rate discussion, and Amrit GPS resources]3.4.2 Launch Vehicle3.4.3 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 Refinement3.6.1 JPL Midterm Mission Design Presentation Feedback3.6.2 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 Analysis3.6.6 Final Flyby Maneuver and System Disposal (NEW)

4.0 System Design4.1 Baseline Designs

4.1.1 Primary ConOps Baseline Trajectory4.1.2 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 Issues4.6.2 Ethical, Social and Health/Safety Issues4.6.3 Manufacturability, Political and Global Impact Issues

5.0 Summary and Conclusions [corrected...but will need final update]

6.0 Design Critique (all NEW)6.1 Strengths6.2 Weaknesses6.3 Confidence6.4 Alternatives

7.0 References

8.0 AppendicesAppendix I: Primary Resources Reference Information [add anything important]Appendix II: FLARE Team Management[updates needed>>tables and contribution

statements]Appendix III: Subsystem RequirementsAppendix IV: JPL Feedback

List of TablesTable 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 comparisonTable 4: Design selection criteriaTable 5: Tentative ConOps selectionTable 6: Thermal requirements[needs updated to current components]Table 7: Baseline trajectory data, departure and heliocentricTable 8: Baseline trajectory data, flybysTable 9: MELTable 10: EVALTable 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 SourcesFigure 1: Primary ConOpsFigure 2: Secondary ConOpsFigure 3: CSD dispenser deployment setupsFigure 4: Sherpa on payload sectionFigure 5: FLARE Primary ConOps PBSFigure 6: INSPIRE cubesatFigure 7: JPL LMRSTFigure 8: Iris X-band TransponderFigure 9: BlackJack GPS receiverFigure 10: Radio Aurora eXplorerFigure 10: Radio band comparisonFigure 11: Launch system analysisFigure 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 heliocentricFigure 14: Baseline trajectory, flybysFigure 15: Component Selection PBS [needed]Figure 16: Timeline for primary ConOps

Acronyms and Symbols

~ Approximately> Greater thana Semimajor axise Eccentricityi InclinationH Altitude of periapsisφ Geocentric Latitudeλ geocentric longitudeVf Inertial spacecraft velocity at closest approachV_inf Hyperbolic excess velocityΔV_inf Anomalous change in hyperbolic excess velocityDA Deflection angleαi Right ascension of the incoming osculating asymptotic velocity vector

δi Inbound declinationαo Right ascension of the outgoing osculating asymptotic velocity vectorδo Outbound declination

ADCS: Attitude Determination and Control SystemAU: “Astronomical Unit”, Earth’s approximate distance from the SunConOps: Concept of OperationsCSD: Capsulized Satellite DispenserDSN: Deep Space NetworkDV: “Delta-V”, a propulsive maneuver resulting in velocity changeEELV: Evolved Expendable Launch VehicleEM: Earth to MoonEPS: Electrical Power SystemEVAL: Equipement Volume Allocation ListEVE: Earth Venus Earth, order of flybys on trajectoryFLARE: Flyby Anomaly Research EndeavorFOTON:GNSS: Global Navigation Satellite SystemGPS: Global Positioning SystemGN&C: Guidance Navigation and ControlHEO: High Earth OrbitJPL: Jet Propulsion LaboratoryJ#: Gravity term of denoted order (#)LEO: Low Earth OrbitLMRST: Low Mass Radio Science TransponderME: Moon to Earth

MEL: Master Equipment ListNEN: Near Earth NetworkPBS: Product Breakdown StructurePEL: Power Equipment ListRAAN: Right Ascension of the Ascending NodeRAX: Radio Aurora eXplorerS/C: SpacecraftSOI: Sphere of InfluenceSLR: Satellite Laser RangingSSPS: Spaceflight Secondary Payload SystemTBR: To Be ResolvedTPS: Thermal Protectant SystemTRL: Technology Readiness Levelwrt: 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 MESSENGER

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 Mission Motivations

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 Mission Constraints 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 Mission Requirements

[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 Design Development

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 Design Alternatives 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.

3.1.1 Preliminary ConOps 1

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.

3.1.2 Preliminary ConOps 2

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.

3.1.3 Preliminary ConOps 3

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 Design Heritage

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 EarthThe 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 characteristicsThe 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 mentionedBipropellant 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 AnalysisA 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 DesignSee 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 BaselinesThis 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 TrajectoryA 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 Design ChoiceThis 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 OverviewFigure 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 Master Equipment 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.

Table 9: Master Equipment List (MEL) of Selected Components

A 6-Unit CubeSat Structure from Innovative Solutions In Space (ISIS) via cubesatshop.com was chosen as the baseline satellite structure because it provided the necessary volume and functionality needed to house the necessary components of the spacecraft. This structure allowed for a modular design that allowed for vertical and horizontal orientation of printed circuit board (PCB) stacks measuring 94 by 94 mm. The structure is slated for multiple launches in the upcoming 12 months and is natively compatible with other ISIS components, as well as components from ClydeSpace.

The Blue Canyon Tech XACT attitude determination and control system was selected as the control system of choice aboard this spacecraft. Its upper limit of 0.007 degrees for pointing accuracy satisfies the pointing requirements that dictate the direction the spacecraft needs to be pointing in order to send and receive signals. Its onboard star tracker and IMU will provide precise attitude determination while the three reaction wheels are designed to provide a slew rate of 10 degrees per second for an eight kilogram, three unit cubesat.

The Iris Navigation and Telecomm Transponder was chosen because it satisfies the requirement for a transponder that functions on X-band frequencies. The Iris transponder is DSN and NEN compatible and provides full duplex Doppler for navigation.

The OEM628 Triple-Frequency + L-Band GNSS Receiver was chosen as the GPS system aboard the FLARE spacecraft. The triple frequency OEM628 natively works with the GPS, GLONASS, Galileo, and BeiDou constellations. Its horizontal position accuracy using the L1 and L2 bands is down to 1.2 meters while its accuracy using RT-2 (from NovAtel) is 1 centimeter plus 1 ppm. These accuracies aid in satisfying the overall requirements for the FLARE mission.

The Andrews Model 160 High Performance Flight Computer from ISIS via cubesatshop.com was chosen as the flight computer for the FLARE spacecraft. Its dual core, 400 MHz processor, 64 MB of SDRAM, and 2 GB of FLASH will allow for adequate subsystem control. While reprogrammable on-orbit, the Andrews Model 160 can also be configured with Linux Real Time Operating Systems, yielding easy integrability with the subsystem components.

The electrical power system, power distribution module, batteries, and solar panels were all chosen from ClydeSpace due to their native compatibility with the 6-Unit CubeSat Structure from Innovative Solutions In Space. These systems are made to work efficiently together in order to provide power to the various spacecraft subsystem components.

The MPS-120XL CubeSat High-Impulse Adaptable hydrazine propulsion system was chosen because it satisfied the delta-V requirements of the mission. In its relatively small 2 U form factor, the MPS-120XL provides 200 meters per second delta-V for a 6 Unit 10 kg spacecraft. The four, 1 N rocket engines also provide the option for momentum dumping from the reaction wheels in the BCT XACT. Additionally, the low operational voltage reduces the load of the propulsion system on the power supply of the spacecraft.

4.2.3 Equipment Volume Allocation List (EVAL)

Assuming a 96x96 mm base, the heights of several potential components were analyzed in order to ensure space within a 6U CubeSat. The systems are as defined in the MEL. Designing to the worst case scenario in the baseline by using using the components with maximum volume from among a sample of components, the components were determined to occupy 4.21U, as depicted in Table 10. These components ended up being the same components used in the final system. This analysis assumed at 10% volume contingency and a 15% margin. The assumption of a 96x96 mm base was a conservative estimation, as Table 3 reveals the actual component bases to be smaller for the most part. This volume analysis indicates that there will be at least 1.79U remaining for a radiator, with potentially more room to spare.

Table 10: Equipement Volume Allocation List (EVAL) of selected components.

4.2.4 Power Equipment List (PEL)

The equipment selected as seen in Table 11 for the power equipment list is a representative set of components that is expected to be on the spacecraft, but component selection is not final. Even at maximum power usage, the spacecraft still has margin on power available. At the minimum expected power output of forty percent solar coverage, there are still several watts of power available even under maximum consumption at 36 Watts. However, the maximum consumption condition is unlikely to occur for extended periods of time since the reaction wheels and propulsion unit need not operate simultaneously.

Power distribution is also taken care of easily given the small size of the cubesat. Power sourcing can be taken care of with the listed EPS, and additional power management can also be carried out with the use of a power distribution module.

The spacecraft will have battery power available for times when the solar arrays do not provide enough power. This could be due to a poor angle of incidence with the Sun or when the spacecraft is eclipsed by the Earth. The spacecraft will be spending the majority of its time in space away from the Earth so available solar power will not be an issue there. When the

spacecraft makes its flybys of the Earth it will be at high velocity and spend a maximum time of less than an hour in eclipse. A 20 Whr battery will act as a backup power source for the spacecraft. This is enough energy to supply the spacecraft for over an hour at full power significantly longer if the spacecraft needs to be put into a low power mode. The battery is capable of withstanding hard vacuum conditions and has an integrated system for maintaining battery temperature. The battery also has overcurrent and short circuit protection as safe measures.

Table 11. Power Equipment List (PEL) of Selected Components.

The GPS equipment 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 within the GPS network. 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 due to the previously listed qualities this makes GPS unit an excellent choice for this mission.

The X-band transponder is designed for (deep to near Earth) space communication using the X-band frequency in conjunction with NASA’s Deep Space Network. The whole system consists of the transponder, receiver, and exciters. It has low noise sensitivity and deviations with a temperature compensated receiver giving a typical noise figure of 2.1 dB and typical sensitivity of -158dB. It is able to transmit up to 30Mbps of telemetry data.It is built to military specifications and has been designed for radiation resistance, greatly increasing its reliability due to the increased rigor in production to meet the military specification.

The patch antennas that have been selected have heritage in many space programs, including cubesats. These antennas are able to support high data rates and over 10 Watts of transmitted power. There is a lot of supporting data on the gain bounds and coverage statistics for the antennas, allowing their expected performance to be calculated. The patch antenna comes in a number of standard form factors, allowing communication in X-band and GPS frequencies, among other options

(retain as an option, but might only use X-band...)The Lithium UHF radio will be used to communicate at rates up to 9600 bps. This radio is compatible with standard amateur radio stations, which relieves any need for an expensive facility for communication. The radio is able to transmit on frequencies between 130 MHz and 450 MHz. It has adjustable communication parameters and customizable frequenencies, deviations, and packet protocols. The default packet protocol is AX.25. This radio system has also been designed for ease of use with digital command and data interfacing and configurable packet source and destination call signs, making it an excellent choice to reduce the complexity of the mission.

There will also be a deployable UHF antenna system in addition to the Lithium UHF radio and patch antennas. This deployable system contains up to four large tape spring antennas, customizable to up to 55 cm in length. Its frequency range has over 10MHz bandwidth within the frequency range specified. It has flight heritage on several missions since July 2010 and has a dual redundant deployment system.

The biggest consumers of power are the propulsion units, reaction wheels, and X-band transponder. The propulsion units consume the most power at startup when they are in their ignition phase, which is short lived. The reaction wheels also consume significant amounts of power, but those can be controlled independently and optimized for minimum power consumption. The rest of the equipment uses relatively inconsequential amounts of power, so timing their usage is not of great concern.

4.3 Mission Timeline

Current mission timeline constraints depend primarily on the trajectory baseline. While the trajectory for the secondary ConOps is TBR, the primary ConOps has been developed to the extent that the initial departure data and the length of each mission leg have been determined, as shown in Figure 16.

Figure 17: Primary ConOps Timeline

From mid 2015 to mid 2018 further project development will take place. Further pre-phase A and conceptualization will take place during 2015 to mid 2016. Fabrication, testing, and assembly will take place from mid 2016 to early 2018.

5.0 Summary and Conclusion

The Flyby Anomaly Research Endeavor (FLARE) proposed design is an affordable cubesat mission whose goal is to gather applicable trajectory profiles regarding the anomaly. In accomplishing that goal we intend to use high Technology Readiness Level components and redundant and complementary platforms for mission and data assurance. The phenomenological formula isn’t precise and only fits anomalies to an acceptable degree when closest approach took place under 2000 km. The data collected by FLARE will further the analysis started by the investigation which lead to the formula’s development and as well as causal investigations.

The primary Concept of Operations (ConOps) incorporates a heliocentric trajectory where an unpowered Earth flyby should be executed on a six month and year alternating basis. A secondary ConOps incorporates a powered flyby of the moon followed by a single unpowered flyby event (meaning multiple deployed satellites and 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 was 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 with respect to the equator (~60 deg), highly elliptic (~.74) and suitably elevated (apogee altitude ~ 40,000

km) parking orbit. Other assumptions are a 10-15% mass/volume/power contingency and 40% sunlight exposure for static solar arrays.

The top-level 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.

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 (Galileo through NEAR). 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 and/or the dominant explanations (high order gravity terms and the anisotropy of the speed of light) 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. 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 discrepancy). 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.

6.0 Design Critique

The entirety of this section is TBR from feedback regarding this midterm report. This feedback will come from 3 primary sources, Dr. Fowler (UT professor) and JPL correspondents Travis Imken and Damon Landau.

6.1 Strengths6.2 Weaknesses6.3 Confidence6.4 Alternatives

7.0 References[1] Michael M. Nieto and John D. Anderson. “Earth flyby anomalies”, Physics Today. Oct 2009.

[2] Anderson, John D., Campbell, James, K., “Anomalous Orbital Energy Changes Observed during Spacecraft Flybys of Earth”. JPL. March 2008. Web. <http://journals.aps.org/prl/pdf/10.1103/PhysRevLett.100.091102>

[3] Jason Andrews. “Spaceflight Secondary Payload System (SSPS) and SHERPA Tug - A New Business Model for Secondary and Hosted Payloads”, Spaceflight, Inc. 26th Annual AIAA/USU Conference on Small Satellites.

[4] “Canisterized Satellite Dispenser (CSD) Data Sheet”, Planetary Systems Corporation. 21 Jul 2014.

[5] “Space Launch Report: Rokot/Strela”, http://www.spacelaunchreport.com/rokot.html#config. 19 Dec 2014.

[6] Antreasian, P., Guinn, J., “Investigations Into the Unexpected Delta-V Increases During the Earth Gravity Assists of Galileo and NEAR”. JPL. Web.

[7] Operational considerations for CubeSats Beyond Low Earth Orbit, http://kiss.caltech.edu/workshops/smallsat2012b/presentations/lightsey.pdf [accessed 02/16/2015].

[8] Orbital Mechanics, ed. Robert A. Braeunig, http://www.braeunig.us/space/orbmech.htm [accessed 02/16/2015].

[9] ISIS. “CubeSatShop.com”,http://www.cubesatshop.com/.

[10] Blue Canyon Technologies. “Products”,http://bluecanyontech.com/products.

[11] SkyFox Labs. “piNAV-L1/FM”,http://www.skyfoxlabs.com/products/detail/1.

[12] Clyde Space. “CubeSat Lab”,http://www.clyde-space.com/cubesat_shop.

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[16] E. Kahr1, O. Montenbruck, K. O’Keefe1, S. Skone, J. Urbanek, L. Bradbury, P. Fenton. “GPS Tracking On a Nanosatellite – The CANX-2 Flight Experience”, 8th International ESA Conference on Guidance, Navigation & Control Systems. Czech Republic, 5-10 June 2011.

[17] Jessica Arlas, Sara Spangelo. “GPS Results for the Radio Aurora Explorer II CubeSat Mission”, American Institute of Aeronautics and Astronautics.

[18] Oliver Montenbruck, Remco Kroes. “In-flight performance analysisof the CHAMP BlackJackGPS Receiver”, GPS Solutions, 2003.

[19] Jonathan Sauder. “Ultra-Compact Ka-Band Parabolic DeployableAntenna (KaPDA) for Cubesats”, JPL, Icube Sat Workshop, Pasadena, CA. May 2014.

[20] S. W. Asmar and J. W. Armstrong. “Spacecraft Doppler tracking: Noise budget and accuracy achievable in precision radio science observations”, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, California, USA. RADIO SCIENCE, VOL. 40, RS2001, doi:10.1029/2004RS003101, 2005.

[21] Michael Christopher Moreau. “GPS Receiver Architecture for Autonomous Navigation in High Earth Orbits”, The University of Colorado, Department of Aerospace Engineering Sciences, 2001.

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[33] Duncan, Courtney. “Iris CubeSat Compatible DSN Compatible Transponder for Lunar Communication and Navigation … and Beyond “. Jet Propulsion Laboratory, California Institute of Technology. Lunar Cubes #3. Nov 15 2013.

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Lisboa. Lisboa, Portugal. Jan 4, 2012.

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8.0 AppendicesAppendix I: Primary Resources Reference InformationSHERPA relevant information and depictions [3, 25].

Figure: SHERPA configuration [3, 25].

Figure: SHERPA propulsion capabilities [3].

Figure: SHERPA deployment/launch configuration, and payload capabilities [3].

Figure: SHERPA mounted on a primary payload of a Falcon 9 [3, 25].

Satellite GPS Link Budget: Radio Aurora eXplorer (RAX) [43]

Figure: RAX CubeSat, GPS heritage section 3.3.4 and Figure 10, GPS link budget [43].

Capsulized Satellite Dispenser (CSD) Data Sheets [4, 44]

Figure: CSD payload specifications sheet [27].

Table: CSD specifications [4].

Lunar Mission Characteristics: Reference Secondary ConOps section 1.5.2

Anomaly Visualization: MATLAB plots

Doppler Noise Magnitudes courtesy of JPL [20]

Appendix II: FLARE Team Management

Table of Contents

1.0 Team Personnel Strengths1.1 Amritpreet Kang1.2 Jeff Alfaro1.3 Graeme Ramsey1.4 Kyle Chaffin1.5 Anthony Huet

2.0 Leadership Descriptions2.1 Project Manager - Amritpreet Kang2.2 Chief Engineer - Jeffrey Alfaro2.3 Systems Engineer - Graeme Ramsey

3.0 Work Breakdown Structure (WBS)3.1 Project Management3.2 Project Systems Engineering3.3 Mission Assurance3.4 Science3.5 Payload3.6 Flight System (Spacecraft)3.7 Mission Operations System3.8 Launch System3.9 Ground System

4.0 Organization Chart and Initial Personnel Assignments5.0 Project Timeline6.0 Cost Estimate7.0 Team Member Contribution Statements7.1 Amritpreet Kang7.2 Jeffrey Alfaro7.3 Graeme Ramsey7.4 Kyle Chaffin7.5 Anthony Huet

1.0 Team Personnel Strengths

The technical and management strengths of each member of the FLARE team are outlined in this section in order to serve as a point of reference for task assignment throughout the development of the mission.

1.1 Amritpreet Kang

Amritpreet Kang is an undergraduate researcher at the Autonomous Guidance Navigation and Control Laboratory (Auto GN&C Lab) at the University of Texas at Austin. He brings technical experience to the FLARE team through his creation of a single marker tracking algorithm at the Auto GN&C Lab. Amritpreet also has extensive management experience as an officer and high power team lead of the Longhorn Rocketry Association. As an amateur high power rocketeer, he has led a team of engineers to successfully integrate timers and altimeters onboard a high powered rocket in order to pressurize and ignite four auxiliary motors simultaneously after takeoff. He brings additional management experience as the treasurer of the Longhorn Rocketry Association (LRA) where he successfully increased the budget of the LRA to its greatest in the organization’s history.

1.2 Jeffrey Alfaro

Jeffrey Alfaro is a student research technician at Applied Research Laboratories at the University of Texas’ Pickle Research Campus. He has experience working in an engineering science environment, as well as leadership, management, and organization experience outside of an engineering capacity. His professional experience has also improved his technical and proposal writing skills. He possesses strong and fairly intuitive grasp of engineering and physics first principles and their applications, as shown through his academic record. He is adept at quickly identifying baseline requirements, as well as spotting potential problems with a design or concept early in the development process. Jeffrey’s work at ARL has also provided him with some knowledge of signal processing and its use in precise position analysis, which is germane to the problem tackled by the FLARE team.

1.3 Graeme Ramsey

Graeme Ramsey is a former assistant in the Texas Satellite Lab (TSL) at the University of Texas at Austin. He brings technical experience through his contribution to the assembly of solar cells/panels and integration or cubesat subsystems at the TSL, in addition to his assistance cubesat integration lends to his awareness of subsystems and interfaces. Also performed CAD renderings (they came back for more so it must have been a good graphic) for NASA, and designed a cubesat deployment system lever in CAD. Graeme has also logged over 100 hours as an amateur rocketeer where he has experience in the fabrication and integration of rockets, in addition to the construction of two solid state motors. Proficient with MATLAB, excel, etc., I also have relevant skills/references as far as performing optimization (monte carlo analysis, mass budget, power budget). I have good leadership and communication skills and can easily function as a backup to the PM. As far as work ethic, you’ll find none better.

1.4 Kyle Chaffin

Kyle Chaffin is technically skilled in areas involving programming and computer aided design. He is strong in modeling, using software such as SolidWorks and has been assigned as the CAD specialist on FLARE. He is also proficient in MATLAB and has experience with the Java and Python programming languages. Kyle brings technical experience from the Longhorn Rocketry Association in which he worked with a rocket payload team for 5 months. During this time, he programmed an Arduino microcontroller, using Java, to track a rocket throughout the duration of its flight by recording GPS data pertaining to the rocket’s longitude, latitude, and altitude.

1.5 Anthony Huet

Anthony Huet is a test engineer for Space Exploration Technologies who has extensive knowledge in rocket engine and test site operation. He has worked as a research intern at the Institut National des Sciences Appliquées de Toulouse where he worked to characterize the responses of a composite structure test rig in addition to using ABAQUS to simulate the flexing and torsion forces applied by actuators. As a member of the Longhorn Rocketry Association, he aided in the design and construction of a supersonic parallel stage rocket with a team of six people. Anthony is proficient in MATLAB and SolidWorks and is familiar with LabVIEW, NX (Unigraphics), Photoshop, Creo (Pro/E), and ABAQUS.

2.0 Leadership Position Descriptions

2.1 Project Manager - Amritpreet Kang

The role of the project manager is to provide managerial leadership to the team in addition to representation. The project manager must organize and coordinate with team members to assure the timely completion of work. Additionally, the project manager must oversee the completion of deliverables and how they adhere to the requirements set for them. In order to ensure the successful conceptualization of the mission, team member progress must be monitored and supplemented by the project manager. As the managerial lead for the team, the project manager must assist the chief engineer and systems engineer in assuring mission integrity.

As the project manager, I will work to obtain regular updates from all of my team members in order to keep the progress of the mission’s development on schedule. I will also understand the work loads, personalities, and commitment levels of all members of the team to the best of my abilities, in order to maintain a positive atmosphere amongst the team. I understand that the main challenges associated with being the project manager are the even distribution of workload with respect to time (to avoid procrastination) and making sure the team gets along with each other.

2.2 Chief Engineer - Jeffrey Alfaro

The chief engineer provides technical leadership for the project. This means that the chief engineer is ultimately responsible for the ability of the mission to meet its goals. The chief engineer reviews the concept of operations to ensure that the design is technically feasible. As the design process continues, the chief engineer is responsible for reliability analysis to assure that the design stands a chance of achieving the mission. Maintainability and safety are also the responsibility of the chief engineer, as a mission which fails due to safety challenges or lack of maintenance cannot meet its goals. Finally, the chief engineer must oversee subsystem designs to assure that they will function as needed and perform in a compatible fashion with the system as a whole. To that end, the ability to recognize potential issues with reliability, safety, and feasibility early on is the strong suit of a good chief engineer.

As chief engineer, I, Jeffrey Alfaro, intend to perform my role by reviewing subsystem designs early to assure that they will perform their role in the system as a whole. I will keep a record of factors and choices affecting reliability, maintainability, and safety throughout the design in order to streamline the process of reviewing those metrics as a whole near the end of the design. I also will attempt to predict problems before they become intractable. The primary challenge of the chief engineer’s function is in understanding the interaction of all the components of the system. Doing so requires that I stay up-to-date on all design choices as they occur and regularly compare their effects with the system architecture as a whole.

2.3 Systems Engineer - Graeme Ramsey

These paragraphs summarize my activity leading up to the proposal submission. As the systems engineer of the group it is my responsibility to oversee trade study analyses, maintain project requirements, and be familiar with technical measures such as power, mass, volume and margins. I kept these factors at the front of my mind especially in regards to ConOps and drawing up the PBS. Going forward I intend to oversee or perform the trade studies and oversee the mass/volume and power budget tables.

Cost and risk are another traditional responsibility of the systems engineer, the PM and I shared this responsibility. I took it upon myself to be the liaison between JPL/TSL and our group by collating and communicating our questions and concerns. As a team we shared the responsibilities of defining and refining the mission architecture and creating our baseline design. My focus during these efforts was to keep track of realistic goals in regards to our mass and power budget, and in the early stages sought to narrow our scope to increase our feasibility.

I intend to do all I can to help this project succeed which includes effectively serving as backup to any and all team members where necessary. Promoting effective communication and systematic progress will be an emphasis of mine going forward from the projects start.

3.0 Work Breakdown Structure (WBS) The work breakdown structure follows from the standard work breakdown structure for a JPL mission. The WBS is broken down as a hierarchical, product driven structure that flows hand-in-hand with the product breakdown structure of the mission. The WBS down to level 2

elements is shown in Figure 1. This breakdown includes nine total elements that define the work associated with the formulation, implementation, and operation of the FLARE mission.

Figure 1

3.1 Project Management

The project management element of the WBS is broken down to level four as shown in Figure 2. The project manager is responsible for maintaining the project’s schedule by coordinating regular and impromptu meetings between the team members in addition to assuring the timely submittal of the deliverables. The project manager must also effectively manage the team of engineers, by properly assigning tasks and monitoring the individual progress of each team member.

Figure 2

3.2 Project Systems Engineering

The project systems engineering element of the WBS is further broken down in Figure 3. The systems engineer responsible for this element must manage requirements by developing and maintaining them. The engineer must also oversee the analysis of trade studies, cost, and risk. Finally, the systems engineer must maintain technical measures such as mass and power, in addition to managing subsystem compatibility.

Figure 3

3.3 Mission Assurance

The mission assurance element of the WBS is intended to oversee and guide the mission’s development, in order to maintain the possibility of a successful mission. Figure 4 shows the mission assurance element broken down in further detail. The engineer in charge of mission assurance must maintain the design integrity of the mission by managing the reliability, maintainability, and safety parameters of the mission. This engineer must also determine whether the mission is feasible within the set timeline.

Figure 4

3.4 Science

The science element is broken down to level 3 in Figure 5. The body in charge of the science element of the mission will work to ensure the investigation of the phenomenological formula developed in order to investigate the hyperbolic flyby anomaly. This body will also analyze data from the mission, in order to confirm the anomaly and update the mission.

Figure 5

3.5 Payload

The payload element is broken down to level 4 in Figure 6. This element dictates the management of onboard instruments through the selection and maintenance of electronics throughout the mission.

Figure 6

3.6 Flight System (Spacecraft)

The flight system element breakdown is shown alongside a product breakdown structure in Figure 7. This element outlines the development, fabrication, testing, and assembly of various spacecraft components and subsystems.

Figure 7

3.7 Mission Operations System

The mission operations system dictates the upkeep and involvement required through the life of the mission. This element is broken down in Figure 8. The mission operations system assures the maintenance of the spacecraft orbit and the management of orbital maneuvers through the calculation and execution of deep space burns.

Figure 8

3.8 Launch System

The launch system element outlines the manner in which the spacecraft should reach the desired initial trajectory. This element is broken down in Figure 9 and entails the selection of a sufficient launch vehicle that can integrate the FLARE spacecraft as a secondary payload. Additional work includes the management of the launcher that will allow a six unit CubeSat to begin its mission trajectory.

Figure 9

3.9 Ground System

The ground system element ensures the optimal tracking of the spacecraft. This element is broken down in Figure 10. The ground system ensures that tracking data be collected and stored in order for the analysis of position and velocity data to occur at a later time.

Figure 10

4.0 Organization Chart and Initial Personnel Assignments The organization chart displayed in Figure 11 outlines the structure of the FLARE team with respect to the overall WBS of the mission. Figure 12 shows the initial personnel assignments of the FLARE team as a restructured organization chart.

Figure 11: FLARE Organization Chart[update needed]

Figure 12: FLARE Initial Personnel Assignments[update needed]

5.0 Project Timeline The project timeline details the team meeting and work plan throughout the Spring 2015 semester of the Mission/Spacecraft Design course. It also includes the deliverables that must be met throughout the semester. The project timeline is shown in Table 1, in addition to the various presentations that will be made over the course of the semester. The timeline does not include the meetings that will take place during class time every Monday, Wednesday, and Friday from 10:00 a.m. to 11:00 a.m., in addition to the weekly team e-mail progress reports that will be sent to Dr. Fowler (as deliverables) every Friday.

Table 1: FLARE Project Timeline

6.0 Cost Estimate

The cost estimate in Table 2 details the amount of money to be spent for the semester’s work on the FLARE mission by the FLARE team. This estimate does not include the implementation of the mission. The cost estimate is based upon the amount of hours that each team member is expected to accrue while working on the FLARE mission over the course of one semester. Higher salaries are given to the project manager, chief engineer, and systems engineer due to the increased responsibilities of the positions.

Table 2: FLARE Semester Work Cost

7.0 Team Member Contribution Statements[update needed]7.1 Amritpreet Kang

As the project manager, my main contributions to the proposals were in the management proposal section. I helped in writing the team personnel strengths section for most of the team members, including myself. I created the work breakdown structure to be a hierarchical representation of the work needed to create, implement, and maintain the FLARE mission and spacecraft. I created the organization chart and assigned initial personnel tasks that ensured the complete coverage of the WBS. In order to maintain a schedule that met deadlines and allowed for the timely completion of work, I created a project timeline that included a team work plan and list of deliverables. I also estimated the cost for a semester’s worth of work by the FLARE team.

With regards to technical tasks, I aided in the conceptualization of the baseline mission design and operational concept, although the majority of the development in this area was done by Jeffrey and Graeme.

7.2 Jeffrey Alfaro

In addition to my statements in regards to my personal strengths and my position as Chief Engineer, my preparation of this proposal involved developing the mission scope to a mature point. In addition, I established criteria for the selection of the concept of operations, which were confirmed by the rest of my team. I performed the analysis of each potential design solution with regard to these criteria. However, in addition to the portions of the proposal which I directly wrote, I also was heavily involved in developing the chosen concept of operations. My role so far has been to identify the baseline solutions. That is, I have consistently identified the minimum system architecture that would meet the mission goals and objectives. I also

developed the first cut system requirements, though these have since been passed along to Graeme Ramsey, our Systems Engineer.

Along with the other team members, I have analyzed the problem of the hyperbolic flyby in order to understand what trajectories would be useful for studying it. I also identified areas in which our team’s understanding of engineering principles or knowledge of feasibility was insufficient and therefore helped formulate the areas of critical information need. As chief engineer, I have successfully found flaws in previous iterations of our scope and improved it, which was instrumental in arriving at our current concept of operations.

7.3 Graeme Ramsey

As the systems engineer it made sense for me to write the proposed trade studies section. On top of that I noted the few critical information needs that weren’t listed as a trade study. My other responsibility in the proposal was to compile the first cut project solutions and system hierarchy. I hand drew several ConOps design options, and a few PBSs (fulfilling my preliminary systems leadership) to accommodate the ConOps architecture. Also I served as the primary means of communication with our JPL associates. I collated our questions and concerns for our mentor and receive valuable information in return from both our mentor and other JPL associates. Also I collated our questions and concerns for our colleagues in the TSL, this proved to be another wealth of information that will continue to be so. On the side I have been researching subsystems involved in a project like ours, from comparing propulsion system, to researching cubesat radiation issues and deployment canisters.

In addition to these factors I made a database for resource compiling via Canvas, a university website. On Canvas everything we had researched can be uploaded into appropriate files for ease of access, as well as creating links to all cooperative workspaces (google docs). By mid-semester, I personally have uploaded over 40 resources (most of which have been valuable) to relevant sections in our database, over 15 relevant MATLAB/EXCEL figures, codes and tables, and I started/outlined/been primary writer for 4/5 of our current collaborations/deliverables (technical proposal, requirements, midterm presentation and midterm report).

For the Technical Proposal I started by outlining most of the sections, making the cover page, the table of contents and structure for the overall proposal. I wrote the executive summary, all of sections 4, 5 and 6. These sections could be summized as ConOps description, selection criteria, ConOps evaluation, project solutions and system hierarchy, launch vehicle considerations, cubesat design criteria, trajectory design, PBS and further design work: proposing trade studies and critical information needs. Also I contributed most of the references in section 7. I read over only a few other sections as a proof read.

I wrote the requirements and justifications where appropriate. These include almost 30 mission level requirements and almost 50 system level requirements.

In regards to the midterm presentation, first I created the structure, by outlining what most of the final slides would be. In totality I made 18/27 slides: the title, overview, executive summary, mission drivers, primary and secondary requirements, constraints, baseline design, all

trade study slides (including GPS accuracy slide), impending refinement, critical issues, the end slide and I contributed most of the references in the reference slide.

In regards to the midterm report, first I created the structure, outlining most sections, creating the table of contents and executive summary to guide the rest of the paper. Listing the sections I wrote in order of appearance: selling statements, mission constraints and assumptions, midterm ConOps, primary/secondary ConOps, system constraints, primary requirements (mission and system), system design development, design alternatives development, preliminary ConOps A/B/C, system and subsystem allocation, data acquisition systems, design heritage, INSPIRE cubesat, X/X-band LMRST, Iris X-band Transponder, trade study summary and results, data acquisition systems TS, launch vehicle TS, ConOps A vs. ConOps B TS, critical parameters, midterm design refinement, JPL midterm mission design presentation feedback, LV and launch trajectory details, subsystem component choices, midterm system design, midterm design choice, system and subsystem overview, summary and conclusions, and design critique. In addition I initiated major revisions to the MEL/PEL/EVAL to reflect the specs contained in certain design heritage references and the PBS. I proof read a few other sections.

In regards to the final report, first I created a list of action items and organized the rounds of assignments from that list. I then communicated all question, concerns and requests for information to Dr. Fowler and our JPL correspondents. I read all feedback from the midterm report and coordinated the correction and improvement process. I corrected all the sections listed above in the paragraph regarding the midterm report and added red comments for necessary corrections and where action items were destined (for organization). I created new sections for the results from certain action items. I also expanded the depth of information provided in Appendix I, regarding our primary resources important information. At this point I also updated all the low level requirements to match the refined higher level requirements. I made two requirements traceability matrices, one for high level mission vs. system requirements and the other for high level mission+system vs. low level system requirements. The new sections that I wrote include: Day in the life of FLARE, Requirements Traceability Matrix,

7.4 Kyle Chaffin

My role in this proposal has been to survey the assortment of literature pertaining to flyby anomalies, providing background information on spacecraft which experienced an anomalous change in orbital energy. I established key orbital heritage parameters of said missions for FLARE to duplicate in determining the existence of flyby anomalies. Furthermore, I assessed various common explanations for the existence of flyby anomalies, and discussed the points at which such explanations fall short, thus necessitating further empirical data from a mission like FLARE.

7.5 Anthony Huet

With my writing skills and general knowledge of spaceflight I wrote the introduction and updated and edited the final version of it. I am also available for feasibility on deep space

propulsion due to my industry experience with propulsion systems. I have also been a part of discussions on potential orbits for the cubesats and other preliminary discussions.

Appendix III: Subsystem Requirements[Level . System . Reference](# . # . #) COMMS0.0.0 The satellites shall have a Comms system capable of receiving trajectory correction commands from the DSN upon departure and approach of Earth.Justification: To achieve recurrent flybys trajectory corrections will be necessary. The most efficient place to apply these DVs is just after the flyby and trajectory determination. Ideally this would set the satellite on the proper course, in reality another small correction will be needed upon approach to assure the predetermined flyby. ADCS0.1.0 The satellites attitude shall be determined and stabilized periodically during its heliocentric trajectory and in preparation for any DV maneuver.Justification: This is an obvious necessity due to the need to point prior to thrusting. We also don’t want our satellite to enter an unstable or rapid tumble state so the occasional stabilizing will be necessary. Sensors0.2.0 During hyperbolic flybys of Earth, the satellites trajectory shall be monitored and analyzed to a fine degree in order to quantify the anomaly.Justification: This requirement is paramount, as viable, precise and applicable (to quantifying the anomaly) data are the sole product of this project. Propulsion0.3.0 The propulsion system shall be capable of providing the DV necessary for the trajectory corrections inherent in setting up repeat flybys.Justification: Multiple flybys are essential so as to maximize the data return of this endeavor. Structure0.4.0 The satellites will be of a standard nanosat size.Justification: This will facilitate the launch vehicle/satellite dispenser and our budget constraints. TPS0.5.0 A thermal protection system shall vent excess heat into space.Justification: 0.5.1 Passive thermal systems shall distribute heat throughout the satellites.Justification: Power0.6.0 The power system will be capable of providing short term power to all subsystems and sustainable power to the COMMs/GNSS system. C&DH0.7.0 The satellites will have hardware and software to manage all subsystems and store any necessary data.Justification: Interfaces0.8.0 Structural interfaces shall integrate the subsystems to the satellites and the satellites to the launch system.Justification: 0.8.1 Electrical interfaces shall integrate the subsystems to each other.

Justification: COMMS1.0.0 The satellites communications system shall also function as a velocity sensor in conjunction with NEN/DSN during each flyby.Justification: This is the same manner in which the heritage missions initially noticed and quantified the anomaly as an anomalous Doppler shift in the satellites output Comms. It will also serve to conserve space in our propellant dominate cubesats. ACDS1.1.0 The satellites’ attitude shall be stabilized prior to each flyby and will no propulsive maneuvers will take place until after the flyby phase.Justification: 1.1.1 The satellites’ attitude will be intermittently corrected during their heliocentric trajectory.Justification: 1.1.2 The pointing accuracy of the ACDS system will be on the order of a tenth of a degree to facilitate accurate trajectory corrections.Justification: Sensors1.2.0 The [satellites’] data post-processing shall achieve an accuracy of velocity data on the order of 0.1 mm/s.Justification: This accuracy was achieved by several of the heritage missions in which the anomaly cropped up. An analysis of Doppler residuals gives an accurate estimation of trajectory and a Doppler shift was the first indication that the anomaly existed at all. 1.2.1 The satellites’ GNSS/GPS receiver shall gather position data during the (less observable) near periapse phase of each flyby.Justification: Propulsion1.3.0 The satellites will be able to escape Earth’s orbit without using their onboard propulsion.Justification: By using a Sherpa assist from an eccentric orbit (e=0.74) an exit trajectory is possible with minimal DV needed directly from the cubesat. The SHERPA 2200 can provide nearly 2600 m/s DV with a small (30Kg) payload. Dictating this limit is essential in order to maintain trajectory correction capabilities over the lifetime of the cubesat. 1.3.1 The projected heliocentric trajectory corrections will not exceed 80% of the satellites’ onboard (after exit maneuver) propulsion.Justification: Structure1.4.0 The structure of the satellites will be a standard 6u cubesat configuration. TPS1.5.0 The thermal protectant system shall consist of passive systems and a radiator.Power1.6.0 The power system shall consist of solar arrays on the surface of the cubesats to supply power, a battery to store power and the necessary wiring.C&DH1.7.0 The C&DH system will be capable of carrying out commands issued from ground stations relayed by the DSN.1.7.1 The C&DH system shall manage collection of solar power and battery charging. 1.7.2 The C&DH system shall gather information relevant to subsystem performance.

1.7.3 The C&DH system shall perform autonomous attitude monitoring and adjustments. Interfaces1.8.0 Lines and valves will serve as interfaces for the propulsion system. 1.8.1 Wiring and a power distribution module will facilitate the distribution of power to all subsystems and components. 1.8.2 Wire harnesses will interface between solar panels and the battery.1.8.3 Wiring will provide an interface from the C&DH system to all satellite sensors and actuators. 1.8.4 A separation electrical connector and tabs will interface the satellites with the CSD. 1.8.5 The CSD will be vertically mounted to the exit assist vehicle (SHERPA). 1.8.6 The exit assist vehicle will be mounted to as a secondary payload in the launch vehicle. COMMs2.0.0 An X-band radio transponder shall function as the satellites’ COMMs system. ACDS2.1.0 Reaction wheels shall perform any required attitude adjustments. 2.1.1 A star sensor in conjunction with MEMs gyros shall perform attitude determination. Sensors2.2.0 The X-band radio shall emit signals during each flyby to verify the Doppler shift associated with the anomaly. 2.2.1 The GPS system will consist of a dual frequency GPS receiver and antenna. 2.2.2 The GPS system shall export position and velocity data to ground stations for post-processing to gain the required accuracy. Propulsion2.3.0 A hydrazine motor taking up approximately half of our cubesats’ volume shall satisfy our propulsion requirements by providing ~200 m/s of DV capability. Structure2.4.0 The satellites structure will adhere to the Planetary Systems payload specifications in reference to compatibility with a 6u (CSD) canisterized satellite dispenser. TPS2.5.0 A surface coating will serve as a passive thermal system.2.5.1 If necessary, a heater shall provide heat to the battery and any other subsystem with heating needs.Power2.6.0 The battery will be of high energy density (~150Whr/Kg) capable of at least 10 Whr (power needs TBR).2.6.1 The solar cells will have an efficiency approaching 30%. 2.6.2 The solar panels shall have temperature sensors, reverse bias protection diodes and harness connectors. C&DH2.7.0 The C&DH system will consist of a space flight computer, flight software and a solid state recorder. Interfaces2.8.0 Details TBR from particular component specifications.

Appendix IV: JPL feedbackJPL sticky note feedbackProgrammaticsWhy does NASA care about this?How does it affect cost, risk and/or scheduleHosting payload to help determine anomaly cause? >> cost-share ConOpsBenefits of COPS A compared to CONOPS B

Can Sherpa provide the dVCONOPS B: what is dV needed @ moon“Macon” Tandem-knowledge of relative position or formation flying?3-body/low-energy orbit? Or resonant orbit? (drawing of orbit around earth/L2)How do we calibrate all aspects of forces on the s/c e.g. SRP, thermal, outgassing, etc…How much separation (time+lat,long) do we need between the two s/c to show “repeatability” in presence of suspected perturbations (eg high-order G.F)Why do we assume we need e>1 for this experiment to work?What other effects are important? >>(day/night flyby)>>(atmo.? [balloons])How long around Earth flyby does the s/c have to be “quiet” to get good data“Randi” Trade Space Slide: to me SOI=Saturn orbit insertion. What does your SOI mean?SLR (Satellite Laser Ranging): should be considered as additional, very-high precision, independent OD system. It is passive on s/c side; just need reflector.

Baseline/Tradespace/Subsystem“Macon” rad-hardened or redundant systems?Spin stabilization? TypesWhy Ka-band?Multiple spacecraft >> hardware, etc. discrepancies?From what range of parking orbits can you departNEAR has data gap at perigee due to DSN slew rate. Fill the gapShould look at (illegible word-“Rou’D”) data rate…can deployed/HGA be avoided?Look into surrey space systems GPS receivers & propulsion systemsACS-can reaction wheels last from SOI entry to exit w/o desats?Small force modeling & calibrationTransceiver vs. transponder. Need a transponder“Macon” UHF/VHF but Ka/S-band antenna?Power/mass for CONOPS B? (3u)“Randi” Power list does not have a star tracker on it.

Problems to UnderstandOut gassing perturbationsJuno: 10X10>>50X50 6-7mm/s dV @ periapsis.

Travis Imken FeedbackFLARE A-Team SessionMarch 16, 2015 9a – 12:15pJPL Attendees:Jackie Green, Randii Wessen, Bill Frasier, Damon Landau, Jeff Stuart, Macon Vining, John Elliott, Eric Gustafson, Melissa VickStudent Team:

Kyle Chaffin, Jeff Alfaro, Amripreet Kang, Anthony Huet, Graeme RamseyMeeting Notes:9:00a – 9:20o Jackie Green introductiono Team introduction and inspirational figures introduction9:20 – 9:40o A Team and JPL introduction9:40 – 11:00o Student presentationsDiscussion on the purpose of the mission: acquire more data pointsRoom discussion on the causes of the flyby anomalyIs the impact based on the shape of the Earth? Could it be the higher order J terms?o Mission DriversCould this mission be done with a highly elliptical orbit vs. a heliocentric orbit?o RequirementsThe “system” will be a combination of ground and spacecraft level components to measure the effecto Secondary RequirementsThere are enormous possibilities with GPS – talk to Bill Frasier to get some more helpo ConstraintsWill use SHERPA for transit. Not included in the costsCost is flexible, may just include hardware for this studyWhat kind of orbits are preferred? There are some 2018 launches that are candidates. There are some possibilities but nothing has been identified so far in this studyo ConOps AWhy are the trajectory correction maneuvers done at the edge of Earth’s sphere of influence?o ConOps BWhat are the deltas if the CubeSats were on a rocket already headed for the moon? Do you still need SHERPA? Maybe. This is a newer idea.Clarification on the ConOps. A is two flybys with a pair, B is one flyby of two pairs in different orbitsWhat is the time between deployment and flyby? Quickly, on the order of days.Eric Gustafson: Outgassing causes accelerations that may be a concern, takes a week or two. Accelerations are on the order that may be a concern.Bill Frasier: From TOPEX, thermal radiation models of all of the surfaces are also an issue. It takes years, decades to model every little thing that could affect. Calibration for this mission ConOps would be difficult.Moon is used as a plane changeWhat kind of separation is desired between the two vehicles to be “repeatable”o Baseline designPrimary customer may be concerned with hydrazine

Need to add in some more sensors on the ADCS?What if the spacecraft was spin stabilized? Might actually add more data to the Doppler for tracking.There is a trade off on spin stabilized vs. 3 axis stabilizationMay interfere with GPS data readings. Putting the antenna on the spin axis will affect where the signal enters the antenna. Would require antenna calibrationHow much are the missions going to try to replicate the previous missions?o Baseline Trajectoryo MELThere is no other payload other than telecomm.Why are there batteries? May be needed in Earth flyby.o Critical IssuesThere are inconsistencies between the MEL, PEL, VEL, and Thermal on the component selection.11:10 – 11:25o Parking lot issuesProgrammaticsCost, risk, scheduleCost sharing with partners with other instruments.Why does NASA care about this?ConOpsCould satellite laser ranging be used? Corner cube reflector (Goddard, UT) would give you an independent and high precision tracking method. Passive on the spacecraft.Compare ConOps A to ConOps BCan Sherpa provide the DV, what is needed at the moon?Do they spacecraft have to perfectly formation fly, or just knowledge.Could spacecraft have super-high accuracy ranging like GRACE? This is just a consideration.Three-body flyby? Talk to Jeff.How do we calibrate all aspects of the spacecraft SRP, outgassing.Why do we need an eccentricity <1 for the experiment to work?What temporal affects matter?How long before or after does the spacecraft need to be “quiet”.SOI means sphere of influenceBaselineConsider rad hard vs. redundancyStabilization methodsDiscrepancies between multiple spacecraft. This relates to vendor variations within parts, characterization, assembly.DSN slew rate at perigeeWhat is required data rate and can HGA be avoided?Are wheels large enough to absorb momentum during flyby without requiring a desaturation?

Transceiver vs. Transponder. Have you identified any CubeSat transponders.Need power/mass for ConOps B.From what range of parking orbits can you depart?Problems to understandOutgassing perturbationsJuno data showed no flyby anomaly in 50 x 50 field. Talk to Eric Gustafson more.11:30 – 11:55o GPS optionsBill Whittaker knows of some GPS optionsConsider the Foton (UT/Cornell)https://www.google.com/search?q=foton+GPS&oq=foton+GPS&aqs=chrome..69i57j0l3.1579j0j4&sourceid=chrome&es_sm=122&ie=UTF-8o What are other ways to frame the problemCould we look at changes in period, etc.Unmodeled changes can add and subtract onto each otherWhy did Messenger and Rosetta not have a flyby error?Elliptical orbitIf we don’t see the anomaly, is it still a valuable mission?Very few Earth orbiters care about mm level of precision? Perhaps those missions don’t care? Lunar and solar orbits in a Molinya orbit may drive you back into the Earth.Could the mission go from hyperbolic elliptical or the other way? These are critical events, which may have not been done on CubeSats before.o Consider making the spacecraft more passive as the ConOps gets more complicatedo CHANDRA may have a highly elliptical orbit, but has a high perigeeo Could the formula be recast to work for highly elliptical orbits.o Look at Lat/Long during periapses and figure out where the S/C is.o How does the light mass of the spacecraft affect the flyby anomaly. Does the mass affect the equation?o Error stackupGround, atmosphere, etc.o Could the mission evolve from elliptical to a hyperbolic over time? This would help the ConOps, difficulties, etc.o Onera makes GRACE accelerometers for CubeSats.11:55 – 12:10o Action ItemsDamon will send papersPrepare the big picture and “why we care”Figure out how to get more dataMore S/c or use other s/c already out thereBalloons

LasersGet information from Sham and TomasEric: Simulation with varying degree and order of gravity fieldsJeff: Look at Surrey Space Systems for hardware, look for PIs that may have proposed explanations for this and look at a cost share and hardware options.Bill: Work on the science traceability for the mission. Tie it back to the observables.Damon: How accurate can we get? How do we knock down the errors?Macon: From a systems perspective, nail down the data and how you are getting it back to Earth?Travis: Look at simpler, passive spacecraft.Randii: Look at 10 x 10 or 50 x 50.John: Look into the overall mission cost.

Damon Landau FeedbackAction Items + Recommendations

•Prepare “big picture”, why NASA cares, what benefits

•Explore “more data is better” and sell it

–More s/c, other s/c already out there

–Balloons, laser, changes in orbit using cubesat prop

•Simulation with degree + order of grav fields

•Hardware – look at Surrey

•Are there PIs who have proposed this & can you cost share

–What would they do

•Create a bridge between science and engineering

–Requirements flow down

–Be critical, tracking data

•What does a null result mean

•How accurate do you need to be, how much can you knock down error

•Systems perspective – data that you will transmit & telecom flow down

•How can you collect the data in the most simple manner

–Reduce risk by conops that get more simple over time

•Consider analysis of alternatives – ok to push back

•Look at “if there IS really a problem here” (10x10, 50x50 gravity)

•Understand mission cost, overall cost – Sherpa, 2 yr of ops

Alternate ConOps Discussion

•DV modeling errors

–Unmodeled changes in error

–Errors stacking up (+’s & -’s)

–Why Messenger & Rosetta didn’t show anomaly

•2&3 flyby powered?

•Elliptical vs Hyperbolic – Does it matter?

–Elliptical – do we care?

–Lunar & solar perturbations really matter

•Put more on ground if possible, e.g. laser ranging

–1U shriek? Cubesat

–Minimize critical events (elliptical orbits come back)

•Learn more about other s/c, Check out s/c on elliptical (e.g. Chandra)

•Change Coordinate system variable

•Cubesats super light compared to other examples

–How does it impact anomaly behavior

–Atmospheric density issues

–Use balloons to gain confirmation of environments

•Grace ONERA accelerometer for cubesat, calculate would it be observable

Outstanding Refinements

•LV + Launch traj details

•Build Earth SOI calcs

•Subsystem component choices

•CAD model analysis

•Final flyby maneuver + system disposalCritical Issues

•Re-evaluate choice based on empirical trade study

•Radiation exposure

•Vibration during launch

•Thermal requirements

•Tracking ability during flyby