flare final report

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

Executive Summary

1.0 Introduction1.1 Heritage

1.1.1 Initial Observations1.1.2 Heritage Missions 1.1.3 Phenomenological Formula

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 1.3.4 Lorentz Accelerations1.3.5 Anisotropy of the Speed of Light 1.3.6 Perturbing Force Error1.3.7 Modeling Error1.3.8 JUNO Findings: Higher Order Gravity Terms

1.4 Mission Constraints and Assumptions 1.5 Report Preview

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 Requirements2.2.2 System Requirements2.2.3 Requirements Traceability Matrix

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

3.1.4 Primary ConOps3.1.5 Secondary ConOps

3.2 Selection of ConOps3.3 System and Subsystems Allocation3.4 Design Heritage

3.4.1 INSPIRE CubeSat3.4.2 X/X-band LMRST 3.4.3 Iris X-band Transponder3.4.4 GPS/GNSS Receivers Overview3.4.5 Satellite Laser Ranging

3.5 Trade Study Summary and Results3.5.1 Data Acquisition Systems Capabilities3.5.2 Launch Vehicle3.5.3 Trajectory3.5.4 Ground Station Tracking3.5.5 Trajectory Separation3.5.6 Propulsion 3.5.7 Prospective Modeling Analysis

3.6 Critical Parameters

4.0 System Design4.1 Baseline Mission Designs

4.1.1 Primary ConOps A Baseline Trajectory4.1.2 Secondary ConOps B Baseline Trajectory4.1.3 Velocity Maneuver Budget4.1.4 Post-Launch and Deployment Details4.1.5 Day in the Life of FLARE

4.2 Satellite Design Choices4.2.1 System and Subsystem Overview4.2.2 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

4.3 Mission Timeline and Schedule4.4 Cost Analysis 4.5 Risk Analysis 4.6 Economics, Environmental and Sustainability Issues4.7 Ethical, Social and Health/Safety Issues4.8 Manufacturability, Political and Global Impact Issues

5.0 Design Critique 5.1 Strengths5.2 Weaknesses5.3 Confidence5.4 Alternatives5.5 Remaining Design Refinements

5.5.1 CAD Model for Analysis5.5.2 Trajectory Refinement5.5.3 Comms Link Budget

6.0 Summary and Conclusions

7.0 References7.1 Image References

8.0 AppendicesAppendix I: Primary Resources Reference InformationAppendix II: FLARE Team ManagementAppendix III: Subsystem RequirementsAppendix IV: JPL Feedback

List of TablesTable 1: Flyby orbital parameters of heritage missionsTable 2: Heritage missions navigationTable 3: Primary requirements traceability matrixTable 4: LMRST comms link budgetTable 5: Noise from various LEO benchmark testsTable 6: Steady state GPS navigation errorsTable 7: Visibility and Slew Rate for potential tracking systemsTable 8: Thermal requirementsTable 9: Baseline trajectory data, ConOps A departure and heliocentricTable 10: Baseline trajectory data, ConOps A flybysTable 11: Baseline trajectory data, ConOps B flybyTable 12: DeltaV BudgetTable 13: Design selection criteriaTable 14: MEL - Master Equipment ListTable 15: EVAL - Equipment Volume Evaluation ListTable 16: PEL - Power Equipment List - NominalTable 17: Power Production at 40% and 70% OutputTable 18: Orbital correction maneuver powerTable 19: Desaturation maneuver powerTable 20: Flyby powerTable 21: IRIS Comms Link BudgetTable 22: Component-wise cost estimation for one 6U cubesatTable 23: Phase D through F cost estimation for two 6U CubeSatsTable 24: Risk Register for the Spacecraft

List of FiguresFigure 1: Magnitude of Potential Error SourcesFigure 2: Simulated Doppler residuals from 7 mm/s anomalyFigure 3: JUNO Doppler postfit residuals reconstruction and deleted dataFigure 4: Position and velocity perturbations from higher order gravity termsFigure 5: CSD dispenser deployment setupsFigure 6: Sherpa on payload sectionFigure 7: Primary ConOpsFigure 8: Secondary ConOpsFigure 9: FLARE Primary ConOps PBSFigure 10: INSPIRE cubesatFigure 11: Downlink rates for INSPIRE using IrisFigure 12: JPL LMRSTFigure 13: Iris X-band TransponderFigure 14: Projected Iris downlink rates for alternate configurations

Figure 15: Downlink rate formulaFigure 16: BlackJack GPS receiverFigure 17: Radio Aurora eXplorerFigure 18: Radio band comparisonFigure 19: Launch system analysisFigure 20: Equatorial and ecliptic planesFigure 21: Ground track for first flyby using ConOps AFigure 22: Ground track for second flyby using ConOps BFigure 23: Radiation shielding using 3D printed materialsFigure 24: ConOps A Departure BaselineFigure 25: ConOps A Heliocentric BaselineFigure 26: ConOps A Flyby 1 BaselineFigure 27: ConOps A Flyby 2 BaselineFigure 28: ConOps A Disposal/Leg 3 BaselineFigure 29: ConOps B BaselineFigure 30: SHERPA mounted on Falcon 9Figure 31: SHERPA deployment from Falcon 9Figure 32: SHERPA rideshare potentialFigure 33: SHERPA 6U CubeSat deployment via a CDSFigure 34: Mission TimelineFigure 35: Risk table and ratings for spacecraft risksFigure 36: Early CAD model

Acronyms and Symbols

~ Approximately< Less than> 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 oscillating asymptotic velocity vector

δi Inbound declinationαo Right ascension of the outgoing oscillating 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 EndeavorGNSS: Global Navigation Satellite SystemGN&C: Guidance Navigation and ControlGPS: Global Positioning SystemGRACE: Gravity Recovery and Climate ExperimentHEO: High Earth OrbitIMU: Inertial Measurement UnitJPL: Jet Propulsion LaboratoryJ#: Gravity term of denoted order (#)LEO: Low Earth Orbit

LMRST: Low Mass Radio Science TransponderMCM: Mid-Course ManeuverME: Moon to EarthMEL: Master Equipment ListNEN: Near Earth NetworkPBS: Product Breakdown StructurePEL: Power Equipment ListRAAN: Right Ascension of the Ascending NodeRAX: Radio Aurora eXplorerSOI: Sphere of InfluenceSLR: Satellite Laser RangingSSPS: Spaceflight Secondary Payload SystemTPS: Thermal Protection SystemTDRSS: Tracking and Data Relay Satellite SystemTPS: 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 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 and 70% exposure for deployed 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 change in velocity for spacecraft to economically reach some distant destinations in the solar system or slow down to reach a captured orbit at inner planets in the solar system. These assists can also be used repeatedly to increase velocity relative to the solar system center of mass and thus significantly decreasing transit time, reducing mission travel time by months or years.

The exact position, angle, and velocity changes experienced by the spacecraft are calculated to great precision. Accurate knowledge of the solar system and physics allows trajectory profiles to be modeled to high precision. Despite this, during some flybys of the Earth the velocity boost that the spacecraft received varied from what was initially modeled. The difference was on the order of millimeters per second, small enough to make little difference to the mission itself, but statistically significant nonetheless. The anomalous DVs were calculated to high precision using Doppler residuals from ground station observations of the flybys. Several explanations for the anomaly have been proposed. To date there is not a sufficient explanation for the cause of this occurrence and thus it remains an anomaly. The proposed mission would be the first of its kind to be launched solely to investigate this anomaly.

1.1 Heritage

This section will provide an overview of the missions which have observed the anomaly and the phenomenological formula associated with the anomaly.

1.1.1 Initial Observations

A flyby anomaly was first detected on December 8, 1990 by JPL’s Galileo I mission engineers who noticed an unexpected frequency increase in the post-encounter radio Doppler data generated by stations of the NASA Deep Space Network as Galileo I flew by Earth to achieve gravity assist [2]. JPL studied this anomalous frequency increase from 1990 - 1993, but no explanation was found [2]. The tracking software was investigated thoroughly along with independent assessments, but no errors were located.

1.1.2 Heritage Missions

While no heritage missions have been dedicated to the study of flyby anomalies, flyby anomalies have been measured indirectly as part of other missions, such as the ones mentioned in Table 1, namely Galileo, NEAR, Cassini, Rosetta, and Messenger. From these missions, we

gather information pertaining to the magnitude of flyby anomalies with respect to various orbital parameters, 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]

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 mm/s. 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.

Table 2: Heritage missions navigation [24-26, 26].

1.1.3 Phenomenological Formula

Phenomenological formulas were developed by Anderson et al. of JPL [2] and Stephen Adler of the Institute for Advanced Study [46] in order to predict changes in hyperbolic excess velocity encountered by spacecraft as they fly by earth. JPL’s model focused on orbital parameters such as incoming and outgoing declinations, while Adler’s model focused on the change in momentum encountered when dark matter particles collide with spacecraft nucleons.

The phenomenological formula developed by JPL, which fits the observed anomaly data, is as follows:

, [1]

The phenomenological formulas developed by Adler are given in equations (2) and (3), in which equation 2 is for the case of an elastic collision between dark matter particles and spacecraft nucleons, and equation 3 is for the case of an inelastic collision.

[2]

[3]

1.2 Mission Motivations

The FLARE mission is devoted to evaluating the existence of a physical phenomenon as the cause of unmodeled energy changes during Earth flybys. Ideally, data gathered by the mission would fill in the near-perigee gap left by most of the heritage missions. Coverage during closest approach could also serve to characterize the anomaly and consequently refine the phenomenological formula. Alternatively, a null result is also informative, in that it increases the likelihood that the anomaly is due to measuring or modeling errors of understood phenomena.

The mission could potentially refine our current understanding of orbital physics. FLARE could result in more precise trajectory propagation modeling. Of particular relevance, the modeling of near-Earth or Earth rendezvousing objects, e.g. asteroids, could be improved. 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. Were near-Earth object orbits to be more accurately propagated, earlier detection of potential hazards would allow action to be taken while small DVs are a viable option.

Other benefits from this project include further advancing the state of the art in regards to the usage of cubesats in deep space missions. It would also serve to further demonstrate and

refine emerging cubesat technologies and techniques in regards to navigation in heliocentric space, including trajectory, attitude, and radiation mitigation. 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 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 most have been ruled out, more data is needed to determine the existence and nature of flyby anomalies. Figure 1 below depicts the magnitudes of some perturbations associated with general satellites in space.

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, Stephen Adler of the Institute for Advanced Study proposed dark matter encircling the Earth [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]. This explanation is therefore inconclusive, though considered less likely than others due to the very high effect predicted. 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

M.E. McCulloch in the Journal of British Interplanetary Society explored modification of inertia as an explanation for the anomaly [30]. A model of modified inertia which used a Hubble-scale Casimir effect could predict anomalous changes in orbital energy on the order of magnitude of the flyby anomalies with the exception of NEAR [30]. However, this explanation lacks experimental testing and empirical data, and is unable to accurately predict a large change in hyperbolic excess velocity as seen in the NEAR spacecraft data.

1.3.3 Special Relativity

Jean Mbelek of Service D’Astrophysique offered special relativity 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

Atchison et al. of Cornell University and Draper Laboratory 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 Perturbing Force Error

According to Antreasian and Guinn of JPL, perturbing forces such as such as relativistic effects, tidal effects, Earth radiation pressure and atmospheric drag can be ruled out as possible sources of error because the imparted acceleration upon the spacecraft is several orders of magnitude less than observed [6].

1.3.6 Modeling Error

Antreasian and Guinn further state that the Galileo I flyby anomaly prompted an investigation of both the navigation software of the Navigation and Flight Mechanics section at JPL and the mathematical models used for deep space navigation [6]. Goddard Space Flight Center and University of Texas Center for Space Research investigated the discrepancy as well, but found no definitive explanation pertaining to the source of the change in hyperbolic excess velocity [6].

1.3.7 Anisotropy of the Speed of Light

Reginald T. Cahill of the School of Chemistry, Physics and Earth Sciences proposed that flyby anomalies are not real and are the result of using an incorrect relationship between the observed Doppler shift and the speed of the spacecraft based on the assumption that the speed of light is isotropic in all frames [44]. Cahill declared this to be a faulty assumption and that the speed of light is only isotropic with respect to a dynamical 3-space and proposed that by taking into account the repeatedly measured light-speed anisotropy, the anomalies are resolved ab initio [44]. Cahill does not however resolve the Pioneer 10/11 anomalies [44].

1.3.8 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]. A simulation of expected Doppler residuals is depicted below in Figure 2. The Doppler residual depicted takes place at closest approach, thus the velocity anomaly is less in magnitude than the excess velocity anomaly. It represents an approximate anomalous excess velocity discrepancy of 6 mm/s. It is important to note that for the JUNO flyby the spin signature of the satellite was preprocessed out of the Doppler residuals, also depicted below.

Figure 2: Simulated Doppler residuals from 7 mm/s anomaly with (left) and without (right) spin signature [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 could produce detectable errors in trajectory propagation comparable to the predicted flyby anomaly [36]. Other possible sources of error such as the three-sigma standard deviation in Earth’s GM and variations in J2 that aren’t well understood in a predictive sense were considered and discredited as explanations, as they were incapable of creating an error that would be strong enough to be easily detected in real-time monitoring [36]. Depicted below in Figure 3 are the actual Doppler residuals recorded from the JUNO flyby, and also deleted data resulting from a burn. While such a burn might invalidate the results, the DV was off track, so that it ought not affect the expected anomaly. However, pointing errors associated with the burn may still be responsible for the lack of an anomaly associated with JUNO, so it does not completely rule out the phenomenological formula on its own.

Figure 3: JUNO Doppler postfit residuals reconstruction (top) and deleted data (bottom) [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 predicted using higher order terms matched the observed trajectory without presenting an anomaly. 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. Unfortunately, such a simulation has yet to be performed, and is recommended as the first step for further efforts to resolve the anomaly.

Depicted below in Figure 4 are the relative velocity and position differences between modeling with a 10X10 gravitational field versus 20X20, 50X50 and 100X100 fields. This figure shows that, indeed, the use of higher order gravity models can resolve the anomaly, the higher order fields approach the anomaly value where 100X100 produces a 6 mm/s (very close to the expected anomaly) difference from 10X10.

Figure 4: Position (top) and velocity (bottom) perturbations incurred by modeling higher order gravity models compared to a 10X10 field [36].

1.4 Mission Constraints and Assumptions

In order to develop a mission capable of observing the flyby anomaly and comparing it to the phenomenological formula, a variety of constraints must be met by the system. In addition, further constraints were imposed by the organization requesting this mission proposal, the Jet Propulsion Laboratory. The constraints are bulleted below followed by rationale.

•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 Deep Space Network (DSN) and Near Earth Network (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 the primary phenomenological formula (1).From observation of the variables involved in the phenomenological formula, it becomes

apparent that in order to produce an anomaly anomaly using currently available methods, a large difference in the cosines of inbound and outbound declinations and large hyperbolic excess velocity are be required, corresponding to an anomaly on the order of mm/s. While the two parameters are coupled, a good first estimate is that the difference in cosines of declinations should be larger than 0.3 and the hyperbolic excess velocity should exceed 1 km/s.

•Mission budget: $5mil before launch associated costs.In order to maximize mission viability it is important that it 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 and SHERPA costs are expanded on in the Cost section (4.4).

•Launch window and parking orbit/exit trajectory characteristics.Regardless of the mission architecture, the constraints applied to the flybys also heavily

constrain the approach to the flyby. The hyperbolic excess velocity requires that the satellite perform maneuvers to achieve it, but the observations require that those DVs not be performed during the flybys themselves, which is where they are most efficient. Further, the approach to the flyby must be in a direction that will produce a perigee far from the equator or poles in order to achieve a high change in declination of the asymptotes. This constraint applies to any viable baseline trajectory (detailed in section 4.1.1). This means that, in all likelihood, 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.

To reduce the fuel consumption needed to achieve the necessary departure trajectory, the initial parking orbit and thus launch trajectory must match the desired heliocentric orbit. The right ascension of the ascending node of the parking orbit or launch must also match the date of departure such that a DV along the orbital trajectory at perigee places the spacecraft on the proper trajectory, if fuel mass is to be optimized. For example, during the equinoxes, the Earth’s equatorial plane is colinear with the ecliptic plane perpendicular to the direction of the Earth’s motion about the sun. This means that the angle between the planes can be added to the inclination of the orbit about the Earth if the RAAN of the orbit is set at 0 or 180 degrees, for autumnal or vernal equinoxes respectively, in order to achieve an ecliptic declination of the outbound asymptote of nearly 90 degrees. At times between equinoxes, the inclination between planes perpendicular to the Earth’s motion is lessened, and thus a greater equatorial inclination of the spacecraft’s orbit is required.

1.5 Report Preview

In order to meet the constraints while simultaneously providing useful data, the mission is best served by first defining the scope, including explicit statements defining the goals, from which requirements may be derived. Afterword, design concepts can be evaluated against the requirements and constraints in order to determine what mission architectures are most likely to succeed at the mission goals. Then, the chosen concepts will be further developed through trade studies and subsequently a baseline design created until a solid preliminary design is arrived at. The preliminary design must then be evaluated to determine what further steps must be taken and the likelihood of mission success. The remainder of this report is concerned with these steps, in the order herein described.

2.0 Driving Statements and Requirements

This section details the FLARE’s 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 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 mission motivations from section 1.2. The system constraints should be referenced to mission constraints from section 1.4. The scope it meant to guide/constrain the project in order to maintain clear and achievable goals and objectives.

2.1.1 NeedSince, 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, it is simpler to attempt to validate the anomaly first, and would assist in later characterizing it. 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 investigate whether the hyperbolic flyby is a real phenomenon, the first step is to test if 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 test whether the anomaly is repeatable, multiple flybys must be performed with nearly the same declination change. To further characterize the anomaly and, potentially, to refine 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 with at least two sets of declination changes, 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. One manner of accomplishing 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 additional flybys with these two spacecraft with varying orbital parameters of the joint flyby. 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. Accurate telemetry and observations near perigee must be collected to mm/s precision and resolution.

2.1.4 Mission

Multiple small satellites will perform flybys of the Earth. 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 propagate the positions of small near-Earth bodies in the solar system, and thus make earlier decisions regarding their use or threat level. The mission also has the potential, if small, 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 damage.

– 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, if the trajectory provides for limited radiation flux.

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. From historical data the magnitude of the midcourse maneuvers (MCM) are assumed to be 10-20 m/s with two per heliocentric leg (total of 40-80 m/s for 2 legs). Our system is prepared for ~150 m/s of total DV which leaves 70-110 m/s for risk contingency and the disposal maneuver. The baseline propellant required is well within the constraints available via hydrazine propulsion, such that the propellant included may be smaller than the upper limit. One major assumption in this regard is that our launch system, the launch vehicle 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 effect 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 eliminate research and development costs and serve to provide mission assurance through proven reliability. Considering cubesats with similar precautions and exposure to radiation in general, the system can be expected to last between 2 and 5 years barring an unexpected rare events.

•Secondary payload considerations.

–Satellites must be compatible with a Planetary Systems Capsulized Satellite Dispenser.

–Satellite mass: 10-15 kg. Max satellite volume: 6u.

Figure 5: 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 , or CSD, depicted in Figure 5. 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 must contact the ejector plate, which provides up to 400N force during launch due to vibration, and optionally

an electronic interface on the +Z or +X/+Y face for the Separation Electrical Connector, which serves as a safe/arm plug [27]. By limiting the size and mass of our CubeSats, the launch associated costs will be minimized. Although we have additional launch system needs, potentially our s/c could allow ride-sharing on the SHERPA, also known as the SSPS as well, and thus the cost would be shared between parties.

•SHERPA must be compatible with the launch vehicle

Figure 6: SHERPA mounted on a primary payload of a LV [25].

The secondary payload considerations serves to maintain the compatibility of the CSD to the SHERPA. The only remaining concern is that the launch assist system, SHERPA, is compatible with the launch vehicle. SHERPA has been designed to the specifications of medium and intermediate class launch vehicles, as depicted in Figure 6, such as Falcon 9, Antares and Evolved Expendable Launch Vehicle, or EELV [25]. The particular launch assist vehicle 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]. Further information is contained in the table in Appendix I.

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 achieved via ride-sharing. The SSPS is assumed to be included in the launch associated costs category with respect to FLARE’s budget. Although it has been considered as a possible concept of operation by NASA JPL, a highly eccentric orbit is not expected to produce a measurable anomaly associated with its closest approach. Finally,

while the anomaly is potentially resolved through the anisotropy of the speed of light and/or accounting for higher order gravity terms, FLARE is operating under the assumption that more data on the anomaly is beneficial to the scientific community in verifying or refuting these claims.

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 become 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 velocity profiles 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 position and velocity of each satellite throughout the flyby to 1 cm and 0.1 mm/s order of 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 if visibility is impeded or if the DSN dishes cannot slew fast enough to track during that high relative speed segment. GPS and/or satellite laser ranging (SLR) monitoring will be able to fill in the gaps of position and velocity data. 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 in regards to the 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, sufficient ground stations to observe the satellite while in the Earth’s sphere of influence, 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 ground stations that can maintain a visual and slew fast enough. 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 monitored by ground stations.

This serves to satisfy our need for velocity measurements over most of each flyby trajectory, thereby identifying if there was a measurable anomaly. Ground station facilities such as DSN or Estrack will be responsible for gathering the velocity profile on the inbound and outbound flyby legs.

{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, specifically X-band, and technologies which have been or are currently 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 requirements traceability matrix is depicted in Table 3. This table serves to visualize how the high level requirements listed in sections 2.2.1 and 2.2.2 are related. Budget, Mission Assurance and Trajectory requirements, which are also important high level requirements, weren’t explicitly listed in those sections and are added for completeness. The primary use of this table is to make sure that the system requirements facilitate the mission requirements. See Appendix III for lower level requirements and the full

traceability matrix relating high level mission/system requirements to lower level system requirements.

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

The most important factors in the the preliminary design of the FLARE system are resolved using the defined scope and requirements previously discussed. These factors include potential concepts of operations (ConOps) and refinement of mission drivers, baseline feasibility studies, including trajectory and product breakdown structures (PBS), data acquisition system determination, accumulation of design heritage understanding. These and other trade studies allow the recognition of critical parameters to drive the remainder of the project.

3.1 Design Alternatives Development

Preliminary brainstorming and research into the flyby anomaly produced several different ConOps scenarios. These ConOps have varying characteristics as to what quality and quantity of data they could potentially return, along with costs and mission timelines. The concepts are titles according to their final evaluation. Therefore, preliminary ConOps are those that were rejected for violating constraints, and ConOps A and B were compared through further trade studies and chosen as primary and secondary architectures.

3.1.1 Preliminary ConOps 1

This scenario involves multiple cubesats, at least 2, on highly eccentric elliptical orbits around Earth. Each satellite would follow a trajectory with perigees at different declinations. It is surmised that the anomaly might be observable in highly elliptic orbits, as consistent with physics. The satellites would perform multiple orbits to determine if the anomaly was notable in captured orbits. After a large number of captured orbits, the satellites would perform a DV maneuver to set themselves on a hyperbolic trajectory and again attempt to measure the anomaly. This option would produce an unknown amount of data, but 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 measured anomaly would be miniscule to non-existent for captured orbits. The phenomenological formula and available data also require that a sufficient change in declination is required between inbound and outbound hyperbolic asymptotes. For a captured orbit, these values are undefined. Instead, the declination of the line of apsides is generally used as an equivalent characteristic. This translates to a plane change for captured orbits, which do not occur in unpowered eccentric orbits. Finally, achieving hyperbolic excess velocity sufficient to measure an anomaly on a final flyby would be impossible within the DV constraints of the individual Cubesats, which would not be assisted by the SHERPA in this ConOps since they would need to be free-flying to make previous observations.


The second scenario involves a single flyby event using a “mothership” and between 6 and 12 3u cubesats. The mothership with docked CubeSats would be perform an EVE boosting trajectory. Upon approach of Earth after Venus rendezvous, the CubeSats would be deployed and perform paired flybys at varying perigee latitudes to demonstrate repeatability for multiple changes in declination. These CubeSats would be uncontrolled ‘dumb’ GPS receivers and X-Band telemetry transmitters. This option would produce a large amount of data across a wide array of parameters, allowing better characterization of the anomaly. The time frame for such a mission would be medium to long, though the cost would be much higher than other ConOps.

With the boost from Venus our satellites would have sufficient excess velocity with respect to Earth such that the predicted anomaly would be on the order of 10 mm/s. This would decrease the needed sensitivity of the ground systems instrumentation or alternatively increase the resolution of the anomaly, aiding to refine the phenomenological formula. Seven data points would be provided in a relatively short time period, including the mothership trajectory profile. Portions of this concept were reproduced in ConOps A, treating the SHERPA as a mothership. However, a mothership capable of an EVE trajectory and communication would necessarily be much larger than SHERPA and incur much greater development costs. This ConOps was therefore rejected due to violation of cost constraints.


The third ConOps scenario is a recurring 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 provide greater heliocentric energy. This microsat would be much more capable than the CubeSats considered in all other ConOps. It would incorporate multiple methods of accurate velocity profile acquisition, and other scientific instrumentation in an attempt to characterize the anomaly and evaluate the proposed causes. This option would produce a low rate of data return, but with very high quality. However, this

mission would incur high cost. More importantly, this architecture’s approach is broad and unfocused. Ultimately, it falls outside the scope and constraints of the mission by attempting to validate several hypotheses at once.

This idea maintains merit if in the event that another mission meets the requirements. That is, if a current mission had planned an unpowered flyby of Earth which would follow a trajectory providing an expected anomaly of measurable magnitude, the velocity profile could be applied to the analysis of the anomaly. JUNO (see section 1.3.8) was one such mission, from which a velocity profile including closest approach was produced after it performed an Earth flyby in 2013.

3.1.4 Primary ConOps A

The primary ConOps, depicted in Figure 7, consists of tandem hyperbolic flybys of Earth by a pair of CubeSats with 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 exit assist vehicle (SHERPA) may also provide an additional velocity profile during the first scheduled 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 SHERPA.

Figure 7: Primary ConOps depiction.1. Launch as a secondary payload into a highly inclined orbit.

The baseline trajectory assumes a launch into a parking orbit with 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 is immediately adopted by NASA or JPL at the conclusion of our study. The trajectory was modeled from its departure from a Molniya parking orbit. Once the launch vehicle deploys its primary payload, the SHERPA 2200 could immediately separate and begin the exit trajectory maneuvers if the launch was nicely matched up with our baseline trajectory. In this scenario SHERPA will deploy after the primary payload and perform small orientation maneuvers to align its orbit in preparation for the departure trajectory. The primary exit DV maneuver will take place at periapse of the parking orbit.2. SHERPA 2200 provides velocity boost for FLARE CubeSats to escape Earth’s infuence.

In performing the above mentioned exit trajectory maneuver, the SHERPA 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 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 aboard the CubeSats will be required only for trajectory corrections, rather than DVs used to significantly change the trajectory. A 10% contingency is added to expected trajectory correction maneuvers from heritage data. 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 achieved by redirecting the CubeSats into Earth’s atmosphere to burn up or into heliocentric space into orbits that will not rendezvous with Earth’s.

3.1.5 Secondary ConOps B

The secondary ConOps, depicted in Figure 8, consists of tandem hyperbolic flybys of Earth by two 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.

Figure 8: 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, similar to a geosynchronous transfer 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. 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 its 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.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 unpowered ME flyby trajectory. The trajectory details can be found in the baseline trajectories later in this report.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.

3.2 ConOps Selection

The preliminary concept of operations were removed from consideration by comparison with the mission constraints, as indicated in their individual descriptions. However, this leaves ConOps A and B in contention. Both concepts meet with the constraints, and are likely to meet the goals of the mission. In order to determine which architecture to recommend, further trade studies were needed, including development of baseline trajectories for each. As will be seen in the relevant sections, ConOps A was selected due to its more efficient use of resources and its remaining available margins for use in further spacecraft development. The baseline trajectories also show that ConOps B is only marginally capable of producing the required data within the mission constraints. This is further discussed following development of the baseline trajectories, which provide an understanding of the distinction.

3.3 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 9. 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 potential substitutes. Hydrazine was selected as the best system after consultation with JPL. JPL advised that hydrazine on a secondary payload was an acceptable risk and not uncommon in recent launches. The largest point of contention is 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 patch antennas along with UHF antennas and radio. The more mature design choices narrowed to 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) pending further trade studies and particular ConOps choice.

Figure 9: FLARE Primary ConOps PBS, orange = primary to mission anomaly data, yellow = datasource, red = in contention.

3.4 System Design Heritage

This section describes the approach used and heritage evaluated to design our system. Dominant heritage is depicted in figures, primarily data acquisition systems and “semi-deep space”, i.e. outside of Earth’s orbit, CubeSat system architecture.

3.4.1 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 10) 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 10: INSPIRE cubesat provided for subsystem design heritage [33].

The downlink rates for INSPIRE are depicted below in Figure 11. This provides a baseline of what to expect our system to achieve or exceed with the latest version of the Iris X-Band transponder. The 62.5 bps line in the figure represents the divide between signals and tones. Tones can still be used to calculate navigation data. [33] Further details about Iris are included in section 3.4.3 below.

Figure 11: Downlink rates for INSPIRE using Iris [33].

3.4.2 X/X-band LMRST

This JPL developed X-band radio transponder demonstrates the components that will go into FLARE’s Communications subsystem (Comms). It is the precursor to the Iris transponder, which is the final Comms design choice, thus it is a good baseline to consider. 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 12 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. An example comms link budget is depicted in Table 4, serves as a good baseline and is directly applicable to the final communication subsystem design choice, the Iris transponder. [34]

Figure 12: X/X LMRST, JPL developed transponder with X/Ka options [34].

Table 4: X-Band LMRST comms link budget [34].

3.4.3 Iris X-band Transponder

The Iris X-band transponder configuration is depicted below in Figure 13. 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 12.75 W of power when in full transponder mode. In receiver mode Iris demands 6.4 W and only using the processor 2.6 W. The patch antennas work on the X-Band spectrum transmitting at 8.4 GHz and receiving at 7.2 GHz. These antennas have a 3 dB bandwidth of ~300 MHz with a peak gain of 5 dB and beamwidth of 80 degrees [48]. In the INSPIRE configuration, the transmitter draws 5 W power and can downlink at 71 kbps at a distance of 1.5 million km. Depending on the range the data rates in regards to communication can vary from 256 kbps to 62.5 bps [53].

Figure 13: Iris X-Band transponder (left) and low gain X-Band patch antenna board (right), courtesy of JPL [33,48].

The newest version of Iris is (as of mid 2015) Iris V2. There are many configurations of Iris and its antennas that achieve various characteristics demanded by diverse missions, an example of various downlink rates from such configurations in depicted below in Figure 14 along with the data rate formula in Figure 15. The FLARE CubeSats will need to function to gather portions of their trajectory profiles from Earth sphere of Influence (~0.0062 AU or ~925,000 km) inward. The CubeSats must also be capable of receiving trajectory correction commands at ~0.01 AU from Earth and the Deep Space Network (DSN). The maximum distance from Earth that the satellites will be is ~0.1 AU on the first leg of the baseline trajectory and a little further for the second leg, however no commands will need to be issued at these far points.

Figure 14: Projected Iris downlink rates for alternate configurations, courtesy of JPL [34,48].

Figure 15: Downlink rate formula [34].

3.4.4 GPS/GNSS Receivers Overview

When examining GPS receivers that would potentially provide post-processed velocity accuracies of millimeters per second, the “BlackJack” GPS Receiver (Figure 16) 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 16: BlackJack GPS Receiver, courtesy of JPL[38].

Figure 17: Radio Aurora eXplorer (RAX) CubeSat [43].

The CubeSat depicted in Figure 17 is the Radio Aurora eXplorer. It serves as a good source of heritage with regards to command and data handling and radiation tolerance in experimental testing [53], and also the electrical power system which has shown years of successful operations. Additionally a GPS comms link budget for RAX, which is located in Appendix I, provides an example of a comms link budget in LEO which is somewhat applicable to our mission. Our mission will gather GPS data during closest approach which is defined, in reference to GPS, as when the satellites the under GNSS constellation altitude of ~20,000 km.

Other GPS models that were considered and then ruled out include the SGR-05U - Space GPS Receiver by Surrey Satellite Technology US LLC, the piNAV-L1/FM (Flight Model) by SkyFox Labs, and the SSBV GPS Receiver by SSBV Aerospace & Technology Group. These GPS models were all ruled out due to their low velocity accuracy, an effect of being designed to only use the L1 band. In the case of the receivers made by Surrey Satellite Technology and SSBV Aerospace & Technology Group, their receivers were limited to 15 cm/s and 25 cm/s velocity accuracy.

Additional receivers that were considered due to their use of multiple frequencies and channels include the OEM series from NovAtel. The NovAtel GPS receivers were highly considered because of their high, centimeter level, position precision and large amount of on-board storage (in some cases up to 4 GB). However, the NovAtel receivers were ruled out because they were not space-ready and only met military standards, in addition to their low TRL.

The Navigator GPS receiver developed by Goddard Space Flight Center was also considered, but ultimately ruled out due to its focus on weak signal acquisition and not on high precision and accuracy.

Ultimately, the FOTON GPS receiver developed by The University of Texas at Austin was determined to by the GPS receiver of choice for the mission. The FOTON receiver is a miniaturized, dual-frequency receiver that was able to achieve centimeter level position accuracy, similar to the level of precision seen with the BlackJack receiver by JPL and certain NovAtel receivers. Various benchmark tests comparing the observable noise from the FOTON to other GPS receivers can be seen in Table 5 below.

Table 5: Noise from various LEO benchmark tests, note PR is pseudorange. [49]

In addition to its high velocity precision, the FOTON receiver utilizes roughly 1 Watt of orbit average power using on-off cycling. This is much lower than the power required from most of the NovAtel receivers that were analyzed.

3.4.5 Satellite Laser Ranging System

Satellite Laser Ranging (SLR) provides near instantaneous range measurement of a satellite with millimeter level precision. This process works by measuring the travel time of light pulses from a ground station to a spacecraft and back. For this to work, the spacecraft must have a special reflector attached to it in order to reflect the light pulses. The ground stations used for this are located across the globe in order to maximize coverage. The network consists of a total of eight stations operating in the United States, Australia, Peru, South Africa, and Tahiti.

Throughout the years SLR measurements have improved orders of magnitude from an initial precision on the order of meters to milimeters. There is currently the next generation SLR2000 ground station under construction, which uses a low energy, photon counting approach with a high repetition rate that represents a quantum technological advancement. This station is capable of providing 24 hour tracking coverage for satellites up to and including GPS altitudes, with a normal point precision of at least 3mm [56,57].

3.5 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.

3.5.1 Data Acquisition Systems Capabilities

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 18 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 (<0.01 AU).

Figure 18: 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 as described in section 3.4.4, and in particular velocity accuracy in regards to post-processing. Listed in Table 6 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 6 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. GPS data collection is supplementary to the trajectory observations provided by Doppler and carrier phase determination observations from ground stations.

Table 6: 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] 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 make it an excellent choice for this mission.

3.5.2 Launch Vehicle

Determining if a smaller launch vehicle like 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 19. 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 19: MATLAB coded Rokot LV analysis, in conjunction with SHERPA 2200, circular orbits, impulse DV.

3.5.3 Trajectory

A preliminary trajectory for ConOps A was found using the patched conics optimization software TRACT. The initial input estimates were determined by constraining the heliocentric legs of the trajectory to integer or half-integer multiples of the Earth’s orbital period for flight-times between rendezvous.

The departure was evaluated from a Molniya parking orbit matching the constraints, i.e., that the departure date and right-ascension of the ascending node were coupled such that the departure took advantage of the Earth’s equatorial tilt with respect to the ecliptic plane in order to achieve heliocentric orbit from a minimally inclined parking orbit. The initial guess for the DV was chosen so that the declination of the outgoing asymptote and hyperbolic excess velocity would result in a heliocentric orbit differing from the Earth’s orbit about the sun only in inclination. If the heliocentric spacecraft velocity is constrained to equal the heliocentric Earth

velocity, then the departure will result only in an inclination change with respect to the sun, producing an orbit that will rendezvous with the Earth after 6 months.

Figure 20: Depiction of equatorial and ecliptic planes effects on departure, and velocity triangles for transition from geocentric to heliocentric frames.

Several permutations of initial guesses using different departure dates and flight-times were required for TRACT to converge. Once a converging solution was discovered, the output was used as an initial guess for a more accurate numerical orbit propagation in NASA’s General Mission Analysis Tool, or GMAT, with additional perturbations. However, GMAT was unable to converge on a DV solution that resulted in the needed flybys using the output from TRACT.

It is suspected that the output from TRACT is insufficient as an initial guess input into GMAT. In most cases, however, patched conics is a close approximation to the final trajectory DVs. Unfortunately, the highly constrained and unusual nature of the trajectory design causes the patched conics approach to be less reliable than usual. This is because the Earth-to-Earth transfer is an unusual orbit in which the ‘third-body perturbations’ caused by the Earth system during the spacecraft’s heliocentric orbit are much larger than typically assumed, since the Earth system remains relatively close to the spacecraft and in the same relative position for the entire heliocentric leg. Therefore, the spacecraft loses a higher proportion of its heliocentric velocity to the perturbation than normally expected. For this reason, the trajectory analysis for ConOps A remains a patched conics approximation, which must be further developed if the mission is to proceed.

The trajectory for ConOps B, however, was developed entirely in GMAT. An iterative approach was taken from an initial patched conics calculation to target a Luna transfer orbit. GMAT’s iterative methods were used to find a DV from a geosynchronous transfer orbit (GTO) that placed the spacecraft into a hyperbolic flyby of the moon.

The moon flyby was evaluated through the use of B-Plane targeting and iterated until a suitable post-flyby trajectory was found. The post-flyby trajectory had a high eccentricity about the Earth, as well as a high inclination. These orbital parameters are conducive to entering a hyperbolic flyby of the Earth with a high declination change in hyperbolic excess velocity, as required by the mission constraints.

From the new Earth orbit, a DV was calculated that would place the spacecraft into a hyperbolic flyby. The DV must occur between the apogee and perigee of the orbit. At apogee, a DV would increase the perigee altitude, rather than resulting in a flyby, and a DV at perigee would not allow observation of an unpowered flyby. DVs closer to apogee are less efficient at increasing hyperbolic excess velocity, but preserve a high change in declination with the right B-Plane targeting. Alternatively, DVs closer to perigee provide a high hyperbolic excess velocity, but reduce the change in declination. Therefore, an eccentric anomaly of 270 degrees was chosen as a compromise for a baseline trajectory, since this is the point at which the spacecraft is traveling parallel to the line of apsides. The resulting trajectory is found in the baseline section.

3.5.4 Ground Station Tracking

Ground station selection was determined by evaluating two parameters: visibility and slew rate. These two parameters together describe the ground station system’s ability to adequately track the spacecraft during flybys. Three systems were evaluated with respect to the parameters. The Deep Space Network, the European Space Agency’s Estrack system, and the TDRSS, or Tracking & Data Relay Satellite System were evaluated, though the last is not a ground station, it offers capabilities that may be needed.

The worst case slew rate for any ground station was calculated to be ~0.35 deg/s at perigee. This assumes the spacecraft flies directly overhead at its closest approach, and that the Earth’s spin is in the same direction as the satellite pass. The nominal visibility and slew rates are shown in Table 7. Fortunately, all but the 70m DSN dish is capable of slewing at a rate needed to observe the flyby, so slew rate is not a major concern.

Table 7: Visibility and Slew Rate for potential tracking systems.

Visibility was determined by comparing the position of ground stations and their visibility to the ground tracks of the expected flybys from the baseline trajectories, as shown in Figures 21 and 22.

Figure 21: Ground track for first flyby using ConOps A.

Figure 22: Ground track for second flyby using ConOps B.

From the table and ground tracks, the DSN does not have sufficient coverage for visibility at low altitudes. However, Estrack’s cooperative network allows it to maintain visibility during closest approach. Since both trajectories pass over the poles, TDRSS provides the best visibility at altitudes lower than 12,000 km, whereas Estrack may be able to track the

spacecraft over the poles, but only by using multiple stations, requiring more patching of multiple observations and thus increasing the error of the measurements.

Ideally, the maximum possible stations will be used to observe the flybys. However, the minimum coverage required is a number of observers necessary to maintain visibility and tracking for ~2.6 days prior to and following perigee of the flyby. This can be accomplished through a combination of TDRSS and Estrack stations working in tandem. The DSN is most useful for communication with the satellite during trajectory correction maneuvers and when the spacecraft is on it’s flyby trajectory, but at a distance exceeding 30,000 km. Ultimately, cooperation between several systems is ideal.

3.5.5 Trajectory SeparationThe trajectory determination for the Primary and Secondary ConOps both are developed

as though only one spacecraft was travelling along the trajectory. However, after separation of the CubeSats from SHERPA, they must be separated by some amount, which may be measured in time or distance. The spacecraft may achieve separation by maneuvering with respect to one another such that one satellite follows the other. If this is the case, the separation distance or time must be determined.

The inner bounds of the separation distance can be considered based on trackability. If we assume a minimum number of ground stations are able to support the mission, such that only one ground station is available to observe a single FLARE satellite at a given moment during closest approach, then the satellites must be separate by a distance that will permit the ground station to track the pass of the first vehicle, then return to a state of readiness to track the following vehicle.

The most difficult time to attempt this is near perigee, since the spacecraft will be moving at a high slew rate with respect to the ground stations. The pass length for the first flyby of ConOps A is about 2597 seconds, assuming that a ground station has 180 degree visibility. The time for the ground stations to then slew back to their initial positions is then 180 degrees multiplied by the slew rate of the antenna. Using the slowest rate capable of tracking the satellites, 0.40 degrees/s, this will take 450 seconds, for a total separation of 3047 seconds. When propagated to the Earth’s sphere of influence, this means that the spacecraft should have a separation of at least 11,651 km during the heliocentric transits.

The outer bounds of the separation depends on the similitude of the flybys. Since ConOps A does not depend on the synodic period of planets, the important parameter for similitude is the direction of the inclination of the equatorial plane to the ecliptic at the time of flyby, since, if trajectory correction maneuvers are performed before the first flyby to ensure similitude, this parameter will affect the ability of the spacecraft to achieve the second flyby.

The rate of change of the direction of inclination of the equatorial plane varies at a rate of ~1 degree/day. For small angle changes, the result is an increased TCM to line up for the second flyby. We can therefore recommend that the separation be minimized to preserve spacecraft fuel, with an outer bound of ~1 day. Increased separation also requires a greater DV to achieve after the separation event, so minimizing this distance has two beneficial effects.

Applying a safety factor of about 2 to the minimum separation, the spacecraft should be ~6000 seconds apart at perigee, or 22,942 km during heliocentric transit.

3.5.6 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. Additional 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 if ConOps A is chosen. Since the thrust maneuvers must be executed while the spacecraft is returning telemetry data, a relatively short amount of time when the vehicle is near the Earth, 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.

The two main electric propulsion systems available for the cubesat are plasma thrusters and ion thrusters. The performance difference of these two systems is rather large, with plasma thrusters having an ISP in the 500-600s range while ion thrusters are capable of ISPs in the thousands. One drawback to electric propulsion systems is that they can have large power and/or voltage requirements, on the order of 80W or 300V, but smaller lower power units are also available.. Another point of consideration is that electric propulsion also produces very low thrust, usually on the order of millinewtons.

A CubeSat Pulse Plasma Thruster with a specific impulse of 590s is capable of providing a delta-V of 83.3m/s with only 10g of propellant and a power draw of 0.5W. The Busek Ion Thruster BIT-1 has an ISP of 2150s and can provide a delta-V of 303.4 m/s with 10g of propellant, at a thrust of 100μN and power usage of 10W. In addition the thruster mass is only 53g, which is significantly lighter than hydrazine thrusters allowing for the synergistic benefits of higher efficiency and less mass for an even greater delta-V.

Bipropellant engines offer high thrust and moderate specific impulse levels. However, bipropellant engines on this size of CubeSat have not been fully developed and integrating a new propellant system is not worth the added risk.

Another option was solar sails. However, these have the lowest TRL of any of the options available. These also have the similar problem as electric propulsion in that they provide very low levels of thrust. In addition, since the flyby must be unpowered in order for the anomaly to

be measurable, the solar sail would have to be detached sometime prior to the flyby event (Earth’s SOI), further complicating the mission.

Monopropellant thrusters have a long heritage in spacecraft applications. They are also a relatively simple system that requires only one propellant. While it is the least efficient method considered, it still provides ample thrust for the spacecraft maneuvers to be completed in a timely manner. Overall these factors made monopropellant thruster stand out as a clear choice for the propulsion system.

3.5.7 Prospective Modeling Analysis

An analysis of the heritage mission trajectories and the modeling associated with them is out of the scope of this report. The modeling analysis would consist of applying higher order gravity models, as the JUNO mission did during its flyby of Earth, to the other heritage missions: Galileo, Cassini, Rosetta, Messenger and NEAR. If the implementation of progressively higher order gravity models more accurately predicts the real trajectory, similar to the results from JUNO, than it would be safe to say the anomaly has been solved. This model should also be applied to the Pioneer velocity anomalies (however this would require greater knowledge of our solar system’s gravity field, to the order of precision we have calculated Earth’s gravity field).

Also, applying the anisotropy of the speed of light to the JUNO flyby trajectory has not been done. Applying this other dominate potential source of the anomaly to the heritage mission with the most detailed coverage would better evaluate its validity as an explanation of the anomaly. This model has been applied to most of the other heritage missions and only has one discrepancy where it didn’t eliminate the anomaly from the heritage mission’s (Messenger) projected trajectory. This method was also applied to the Pioneer 10/11 velocity anomaly without being resolved.

3.6 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 4.1.4 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.

•Tracking ability during the closest-approach phase of each flyby

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 periapse, according to JPL trajectory propagating models 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.

•Radiation mitigation and 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 Low Earth orbit, as provided by the manufacturer specifications [9, 10]. 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. Additionally, for the majority of the first heliocentric leg the CubeSats will be contained in the capsulized satellite dispenser attached to the SHERPA and thus not in direct exposure to the Sun’s radiation. Generally, radiation effects are mitigated by either using space-rated electronic components or by using a shielding material to attenuate radiation on electronic components [54]. Radiation hardened components multiply the price of the system, thus strategic shielding is the best option. Milled aluminum is the traditional material for radiation attenuation, however the emerging technologies of 3D printed (also called additive manufacturing or AM) radiation shielding using polycarbonates is an alternative with great potential. AM provides two techniques to maximize their radiation shielding capabilities by combining material properties: in changing materials to form layers or a merging materials into a hybrid source [54]. A test of such AM materials is depicted below in Figure 23, where the materials impregnated with tungsten performed to the highest shielding capability, and polycarbonate along with ULTEM materials performed well [54].

Figure 23: Radiation shielding using 3D printed materials [54].

• 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.

•Thermal requirements

The operating temperatures of various selected components aboard the spacecraft are given in Table 8. 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 8: Thermal requirements, Primary ConOps system design.

4.0 System DesignSee subsection 3.1 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 the baseline mission design 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 A 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 9, will place the spacecraft into its departure trajectory, resulting in a hyperbolic excess velocity 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 25, will place the spacecraft on a course to rendezvous with the Earth in half a year. The flyby at that time, shown in Figure 26, 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 10 gives the orbital parameters of the flybys and their predicted anomalous energy changes according to the phenomenological formula.

Figure 24: Departure depiction. The red, green, and blue axes are the MJ2000Eq geocentric axes. The yellow vector is the Earth-Sun vector, and the red trajectory is the parking orbit and

departure.

Figure 25: Baseline heliocentric trajectory depiction. The green trajectory is Earth’s orbit about the sun. The red trajectory is Leg 1, and the cyan is Leg 2. The axes are for the Sun ecliptic

frame.

Table 8: Relevant data for baseline departure and heliocentric trajectories. Departure is in MJ2000Eq geocentric frame, where the heliocentric legs are described in the Sun ecliptic frame.

Figure 26: Flyby1 for Conops A. The yellow vector is the Earth-Sun vector. The axes depict the MJ2000Eq geocentric frame.

Figure 27: Flyby2 for ConOps A in the MJ2000Eq geocentric frame. The Earth-Sun vector is depicted in yellow. The perigee position is clearly very polar.

Table 10: Primary ConOps A baseline flyby 1 and 2 relevant data.

After the flybys, without further DV, the spacecraft will continue along Leg 3, as depicted in Figure 28. Leg 3 has a smaller semi-major axis than the Earth’s orbit. Thus, at perihelion, the spacecraft can perform a maneuver to lower its aphelion or perform a plane change, placing the CubeSat’s orbit inside the Earths in such a manner that it will no longer rendezvous with the Earth, eliminating the likelihood of creating orbit debris about Earth.

Figure 28: Leg 3 of ConOps A.

4.1.2 Secondary ConOps Baseline Trajectory

The baseline trajectory for the secondary ConOps B begins from a parking orbit based on a geosynchronous transfer orbit of semi-major axis of 24,000 km and an eccentricity of 0.715. A slight inclination is applied to the parking orbit in order to rendezvous with the moon from the desired direction. The inclination is 12.4 degrees from the equatorial plane.

The spacecraft performs a burn at perigee in order to place it onto a lunar transfer orbit (LTO) and performs a lunar flyby in order to gain geocentric velocity and to perform a plane change on the orbit. From the new geocentric orbit, the SHERPA burns the remainder of its fuel in order to set the spacecraft on a hyperbolic flyby trajectory. The parameters of the flyby are given in Table 10.

Figure 29: MJ2000Eq geocentric depiction of ConOps B trajectory. The red trajectory depicts the initial LTO, flyby, and following geocentric trajectory to the final DV placing the spacecraft into

a hyperbolic flyby.

Table 11: Flyby parameters for ConOps B.

4.1.3 Velocity Maneuver Budget

Velocity maneuver (DV) values were calculated for both trajectories and are depicted in Table 12. The mid-course maneuvers (MCMs) are planned as trajectory correction maneuvers only for ConOps A, to ensure that the satellites remain on the trajectories for flybys. No large scale maneuvers need to be performed in order to create the trajectories outlined in the baseline for ConOps A.

MCM1 is performed after Earth departure, on the way to the first flyby for ConOps A, or the lunar flyby for ConOps B. MCM2 is performed shortly before that first flyby. MCM3 is performed after the first flyby. MCM4 is the final TCM used to assure the 2nd, or in the case of ConOps B, the primary mission, flyby.

Table 12: Velocity maneuver budget for both ConOps, for SHERPA and the individual FLARE satellites. MCM values are effectively contingencies for perturbations to the planned trajectory

for ConOps A.

Table 11 makes it clear that ConOps B exceeds the DV available to the SHERPA spacecraft, while ConOps A leaves a margin of 400 km/s, the difference between the SHERPA’s maximum capability and the available budget for a fully loaded ride-share.

Therefore, ConOps B is likely unsuitable for this reason alone. However, this is only a baseline trajectory and could likely be optimized. Making better use of the lunar flyby by performing a powered flyby would allow a DV savings that may make the mission feasible. However, it would remain marginally so. This is worth bearing in mind when comparing the two ConOps.

4.1.4 Primary/Secondary ConOps Evaluation

Table 13 serves to enumerate the importance of each criterion used to decide whether ConOps A or B should be pursued. 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. Rate of Data Return 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.

Table 13: Design selection criteria and weight, for ConOps (primarily) and system evaluation.

•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.

ConOps A provides anomalies of ~7.5 and -6 mm/s, well within the desired bounds. ConOps B’s baseline, however, produces an expected anomaly of only -1 mm/s, with the same investment of resources. This is only marginally satisfactory and is more likely to be obscured by unexpected events or variations in data error.

•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.

ConOps B would have a lower launch cost but higher component cost, since 4 CubeSats are involved. However, the launch costs of ConOps A may also be mitigated by available ride-sharing on the SHERPA.

•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 5 data points (velocity profiles that have an expected anomaly). The secondary ConOps allows for 5 data points as well. The two ConOps are equal in this respect

•Rate of Data Return (~2 data points per year)This is the least important factor to the success of FLARE. 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 5 data points in 2 years; this gives

it a ratio of 0.25 data points per month. Whereas the secondary ConOps gathers 5 data points in less than 2 months, giving it an approximate ratio of 2.5 data points per month. ConOps B is therefore superior in this metric.

•Mission Assurance

The primary risk of mission failure comes from potential failure of important mission events such as burns or separation procedures. Since ConOps A uses long heliocentric flights, a missed DV is mitigated by the continued opportunity to correct the trajectory, up until the first flyby. Even without TCMs, the flybys are still likely to produce valuable data.

ConOps B provides a much shorter timeframe in which to perform DVs. However, since the burns take place in Earth orbit, a missed burn allows a later opportunity to re-attempt. For the LTO, this may require a month wait. If a powered flyby about Luna is attempted, the opportunity will be much more constrained.

However, a greater division between the two ConOps appears when anticipating separation of the CubeSats from SHERPA. ConOps A allows the separation to occur after MCM1, providing a 6 month period for alternative plans should the separation not go as planned. ConOps B provides a much shorter window, since the separation must occur after the final SHERPA burn. Thus, SHERPA will have little to no fuel remaining for contingency operations should separation not go as planned. Additionally, the timeframe for separation is reduced to hours.

From these characteristics, ConOps B is only superior in the rate of data return, which is the least significant of the criteria. It may be superior in cost. However, ConOps A is not only more feasible, as shown by the DV budget, but also performs better in the primary criteria of data quality.

ConOps A was therefore selected as the primary concept of operations, with ConOps B as a backup concept. ConOps B is useful as a backup as it may prove useful in the event that ConOps A develops problems, such as finding ride-sharing on a Molniya type orbit, or if the time table for observations must be drastically reduced.

4.1.4 Post-Launch and Deployment Details

The general post-launch procedures are depicted below in figures 30 through 33. The SHERPA will be attached to the payload section of a medium or intermediate class launch vehicle (a Falcon 9 is depicted). The SHERPA will be deployed into the parking orbit described in the baseline trajectory section and primary/secondary ConOps sections. Ridesharing includes other microsats and CubeSats on the SHERPA itself and also a primary payload attached to the payload section via SHERPA, all of which will be deployed before our CubeSats. Most of the ridesharing mass will deploy prior to the main DV maneuver into heliocentric space (or the moon’s sphere of influence for the secondary ConOps). Our CubeSats will deploy once SHERPA has performed all the potential DV maneuvers in heliocentric space (or in the moon’s sphere of influence) to put the system in line for the first flyby. The CubeSats will be ejected

from their Capsulized Satellite Dispensers (CSDs) on the SHERPA in the week (or two) leading to the first flyby, giving the CubeSats time to attain the separation (using minimal fuel) necessary to track each satellite throughout closest approach without interruption. Being contained in the SHERPA through most of the first heliocentric leg will minimize CubeSat radiation exposure early on. The SHERPA will also perform this first flyby after which it will perform a disposal maneuver. This is due to SHERPA’s limited lifetime (~1 year), thus it is uncertain if it can be controlled for the second flyby.

Figure 30: SHERPA mounted on a primary payload of a Falcon 9 [25].

Figure 31: SHERPA deployment from Falcon 9 payload section [3].

Figure 32: SHERPA rideshare potential [3].

Figure 33: SHERPA 6U CubeSat deployment via a CDS [4].

4.1.5 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/outbound flyby phase, and 3) closest approach flyby phase. The secondary ConOps would follow 2) and 3) with some additional consideration to the deployment phase after the moon assist.

The heliocentric phase attitude will be such that the CubeSats deployed solar panels are pointing toward the sun intermittently. A spin about the z-axis will allow passive thermal distribution. 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. The general attitude state will be a stable spin or tumble to minimize the usage of reaction wheels to correct gravity and solar pressure perturbations. Particular attitude maneuvers will be tied into mid-course maneuvers (MCMs) used to optimize the trajectory. In order to make dual use of the MCM burn, the reaction wheels will be desaturated at that point. 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/outbound 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. A slow spin about the z-axis during this phase is acceptable. The Doppler signal can be preprocessed to remove the spin signature. During this phase the trajectory profile will be observed by one of more ground stations.

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. Saturation of reaction wheels is a major issue during this phase because propulsive maneuvers are not allowed throughout the flyby. No spin is preferable during this phase because the CubeSats will have to slew relatively quickly to point at the station of interest.

4.2 Design ChoiceThis section will outline FLARE’s 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.

4.2.1 System and Subsystem OverviewThe Product Breakdown Structure, depicted in section 3.3, gives a visual overview of the

subsystem allocations associated with the CubeSat system. The system is comprised of standard subsystems with respect to heritage missions with similar trajectories and requirements. The components considered are described in the following sections (MEL/PEL/EVAL) which serve to evaluate the CubeSat system 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. The Power system is composed of a solar array, battery, electrical power system, and power distribution module. The thermal protectant system 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. Torque rods could be useful if saturation is foreseen as an issue during the “quiet” flybys of Earth, however severe time constraints with respect to when the torque rods can be used may render them as a waste of mas. The structural choices are a 6U shell with interfaces for the CSD (described in section 2.1.5), 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 and X-band patch antennas (4). 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).

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 SLR capable ground stations. 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 14. These components satisfy the requirements outlined in section 2.

Table 14: Master Equipment List (MEL) of Selected Components, superscripted references.

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 Inertial Measurement Unit (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 Deep Space Network (DSN) and Near Earth Network (NEN) compatible and provides full duplex Doppler for navigation.

The FOTON GPS Receiver was chosen as the GPS system aboard the FLARE spacecraft. Its high precision and low power requirements make it an ideal candidate for the FLARE spacecraft (see section 3.4.4 on GPS/GNSS Receivers).

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 change in velocity for a 6U 10 kg CubeSat. 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.

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 low Earth orbit (LEO) missions. NASA (National Aeronautic and Space Administration) and ESA (European Space Agency) 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 space 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 a rad-resistant option, 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

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 15 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 15: Equipment Volume Allocation List (EVAL) of selected components, superscripted references.

4.2.4 Power Equipment List (PEL)

Power equipment lists (PELs) for common situations the FLARE CubeSats would encounter are detailed in this section. In the event the CubeSats draw more power than the solar panels can produce or in the event the spacecraft is eclipsed by the earth, a 20 Whr battery will act as a temporary power source. The BCT XACT was analyzed with respect to two power states. The low power standby mode utilizes 0.85 W of power. At its maximum power usage of 2.83 W, the BCT XACT operates at 5 Hz with two reaction wheels at 600 rpm and one reaction

wheel at 6000 rpm. The IRIS transponder was analyzed at three independent states. At standby, with only the processor functional, the IRIS transponder operates at 2.6 W. With only the receiver on, it operates at 6.4 W. Lastly, the fully operational transponder operates at 12.75 W. The MPS-120 XL operates at under 4 W during startup but only needs less than 1 W during operation. From this information on the propulsion system, it was assumed that the MPS-120 XL needed 4 W at any operational stage and 1 W during standby, in order to assume the maximum power consumption at any given stage.

A power equipment list (PEL) representative of the nominal power usage on the spacecraft along its heliocentric trajectory is shown in Table 16 below. For this analysis, the BCT XACT was set to operate on low power standby mode in order to preserve power. The IRIS transponder was set to operate at 2.6 W, such that its processor stays online in the event of a necessary correction maneuver. Additionally, the propulsion system was set to standby. The margin was calculated by taking the total nominal power production for 40% solar panel output, as described in Table 17, and subtracting the power consumption for the spacecraft bus. This yielded a 50% power margin.

Table 16: Total Nominal Power Usage

Table 17: Power production at 40% and 70% output levels from the onboard solar panels.

In the case of a needed orbital correction maneuver, a modified PEL was constructed as shown in Table 18. For this analysis, the BCT XACT was set to operate at its maximum frequency, while two of its reaction wheels operated at 600 rpm and one of the reaction wheels operated at a maximum speed of 6000 rpm. The IRIS transponder was set to receive signals only, and the MPS-120 XL was set to maximum power. The margin was calculated from a 70% output from the solar panels. This analysis yielded an approximately 32% power margin.

Table 18: Total Orbital Correction Maneuver Power Usage

In order to estimate the power usage during reaction wheel saturation, another PEL was constructed as seen in Table 19. During this maneuver, the BCT XACT was set to the same specifications as for the orbital correction maneuver. The IRIS transponder was set to operate at 2.6 W in order to maintain the operation of its processor. The MPS-120 XL was set to operate at maximum power. Using the 70% solar panel output to calculate the margin, the power margin was calculated to be just over 40% of the total power output of the solar panels. If the 40% solar panel output was used to calculate the margin, a margin of 0.367 W would be yielded from the calculation.

Table 19: Total Desaturation Maneuver Power Usage

In the case of an unpowered flyby, the PEL seen in Table 20 was developed. For the flyby, the BCT XACT was set to full power, as in the previous two cases. The IRIS transponder was set to send and receive signals. The FOTON GPS receiver was set to operate at maximum power in order to receive signals from visible GPS satellites. The MPS-120 XL was turned off in order to preserve the integrity of the data from the flyby. Using the case of 70% solar panel output, a margin just over 3.5%, equivalent to 4.302 W, was calculated.

Table 20: Total Flyby Power Usage

4.2.5 Comms Link Budget and EbNo Analysis

Table 21 below depicts a preliminary Comms link budget using the Iris X-Band transponder at different points during the flyby phase along with the approximate point at which the midcourse maneuver (MCM) will take place. This table will be a useful reference but should be considered a rough estimate. It is important to note the MCM link budget data is included for reference only, commands will only be sent to the spacecraft at that range not from the spacecraft back to Earth.

The specifications for Iris were gathered from JPL reports regarding Iris in particular as well as its implementation via the INSPIRE CubeSat [33,34,35,45,48]. Calculations of pointing loss and atmospheric loss were calculated to be approximately -0.15 dB each. The pointing error was assumed to be ~1 degree with a beamwidth of 80 degrees. The loss due to the atmosphere from water and oxygen was approximated as 0.015 dB/km and the thickness of the atmosphere that causes this loss was approximated as 100 km..

Table 21: Approximate Comms link budget for Iris integrated into FLARE CubeSats.

4.3 Mission TimelineCurrent mission timeline constraints depend primarily on the trajectory baseline. While

the trajectory for the secondary ConOps, the primary ConOps has been developed to the extent that the initial departure data and the length of each mission leg have been determined. The overall schedule for development of the mission is detailed in figure 34 below.

As shown in depicted in the figure, Pre-Phase A (concept studies), Phase A (concept and technology development), and Phase B (preliminary design and technology completion) have been touched upon and detailed in this report. Upon completion of Phase B, Phase C would involve conclusion of the final design and the start of fabrication. Phase D would involve system assembly, integration and testing, and launch. The mission timeline from the end of Phase D through Phases E (operations and sustainment) and F (closeout) is also detailed in the figure below.

Figure 34: Project Development Timeline (top) and Primary ConOps Timeline (bottom).

As detailed in Figure 34, launch is scheduled to occur during April 2018. The trajectory was modeled using a departure date, from the parking orbit, of May 15, 2018. The first and second legs would take half a year and a full year respectively. The mission would then enter Phase F (closeout) with the disposal of the spacecraft and the post mission data analysis of the X-band transmission from the two spacecraft.

4.4 Cost Analysis

Cost analyses were conducted for the the primary concept of operations and its two spacecraft. In order to estimate the cost for the spacecraft bus and payload, the cost of each component was estimated and added. This cost estimate is represented in Table 22 below. In the cost estimate for the IRIS transponder, estimates from various heritage missions were taken into account in order to provide an extremely rough approximation of the cost. Because the MPS-120 XL was still under development as this report was written, its cost was estimated from resources taken from the Delft University of Technology. It was approximated that a 1 Newton monopropellant thruster would cost about $60,000. It was also approximated that a 0.161 cubic meter hydrazine tank would cost $53,400. Including the cost of four, 1 Newton thrusters, one hydrazine tank, and a ten percent contingency, the overall cost of the MPS-120 XL was estimated to be roughly $322,740. Errors in this estimate can be attributed to price changes and inflation over time, in addition to the usage of a much smaller hydrazine tank than the 0.161 cubic meter tank used for the cost approximation.

Table 22: Component-wise cost estimation for one 6U CubeSat.

Using the total CubeSat cost of $809,311.80, which becomes $1,618,623.60 for two CubeSats, the cost for Phases D through F (see section 4.3 Mission Timeline and Schedule) were estimated. In order to estimate the cost of Phases D through F, the Small Spacecraft Cost Model (SSCM) was modified to fit our use. Developed by The Aerospace Corporation in 1996, the SSCM is a parametric cost model that estimates the cost for a first unit, Earth orbiting spacecraft. It is a cost estimating relationship that predicts a collection of non-recurring and recurring costs. The SSCM bases its estimates on 53 satellites with masses mostly under 100 kilograms. Its estimates of the spacecraft bus and payload costs, however, are extremely inaccurate for satellites under 20 kilograms and over 400 kilograms of mass. In order to compensate for the model’s absolute standard error in spacecraft bus estimation of $3,696,000, the CubeSat component cost estimation, from the table above was substituted as the spacecraft bus and payload cost. The component cost estimation was multiplied by two and used as the cost driver for the CERs in the SSCM, in order to obtain the costs for both FLARE CubeSats. The cost for Phases D through F, using a modified SSCM are detailed in Table 23 below.

Table 23: Phase D through F cost estimation using a modified SSCM for two 6U CubeSats.

The final cost of the FLARE mission, not including any launch costs incurred through rideshare from the launch vehicle or the purchase of SHERPA, comes out to roughly $2,419,842.28. This cost includes the component cost of two, 6U FLARE CubeSats, spacecraft integration, assembly and test, program level costs, flight support, and ground equipment. Additionally, it is assumed that the SSCM compensates for the cost of contractor program management and systems engineering costs. Possible sources of error from the use of the modified SSCM may come from the use of cost inputs outside the input range for the model.

The cost of ridesharing with the SHERPA is $995,000 for a 6U CubeSat in a geosynchronous transfer orbit [55]. As the FLARE mission will use two 6U CubeSats, the cost of ridesharing with the SHERPA will be around $1,990,000. However, the actual cost of ridesharing is likely to be higher than this, as the FLARE mission requires an unusual trajectory. Thus, the total cost of the FLARE mission, without the cost of the SHERPA, will be at minimum $4,409,842.28. The purchase and operation cost of the SHERPA system with 500 kg payload capacity is $20-35 million, though this cost could then be recouped through offering tertiary payloads ride-sharing opportunities. However, these payloads would have to match the parking orbits used in the selected ConOps.

4.5 Risk Analysis

In this mission there are no new risks that are not generally present on other missions. The spacecraft exists essentially to broadcast its location, which does not pose any unique risk. There are the inherent risks at launch with any satellite that the launch vehicle will fail and the mission will be a total loss. The only possible mitigation for this is to select the most reliable launcher possible, but cost will likely limit the options available. After deployment there is risk that a component will fail before the mission is completed. This could be an electronics failure or a physical failure like a reaction wheel or engine. A physical failure could occur from poor assembly of a component, but this is hard to guard against as the vendor must be trusted to produce a high quality component. The high radiation environment of space also poses an ever-

present risk to the electronics. The main preventative measure is using radiation hardened components to reduce the chance of cosmic radiation interrupting the electronics. There are no special mechanism operations to worry about failing, the largest risk is just the potential for random failure in one of the components.

Table 24: Risk Register for the Spacecraft

Table 24 is the risk register for the mission that lists the various risks the spacecraft will face during the mission. This includes a brief description of what the cause of the risk is, the severity of the effect of the event on the mission if it were to occur, and the likelihood of this event occurring. There is also an overall “risk rating” assigned for each risk that evaluates the total risk as a combination of the severity and probability. Finally, potential mitigations for each risk are listed to provide measures that can be taken to reduce the danger that each risk poses by either reducing the severity or probability of the risk happening.

Figure 35: Risk table and ratings for spacecraft risks

The risk table is shown in Figure 35 to help organize the risks present in the mission. Based on a combination of impact and probability the risk is assigned a “risk rating” of low, medium, high, or very high. Increasing probability and/or severity results in a higher risk rating. The item number for each risk is placed in its corresponding square, which presents the various risks in a very useful visual way. This mission has five medium risks, and two high risks. The two high risks are vibration damage at launch and an electronics failure. The risk table helps with easily identifying the highest risk items, so that they may be evaluated and cautions taken to reduce their overall risk rating.

4.6 Economics, Environmental and Sustainability Issues

The economic impact of this mission is extremely low relative to previously flown and current deep space mission. However, the potential for a low science return gives rise to the concern that the mission may not be worth the money spent to develop it. This, however may be remedied by the potential benefits this mission could deliver. If the FLARE mission delivers results that confirm or help to characterize the hyperbolic flyby anomaly, it would serve as an aid to all future missions that involve Earth flybys. This would allow for more accurate mission design. Additionally, this mission could serve as an aid to further develop deep space CubeSats and their applications. Because there are very few components and missions currently geared towards designing CubeSats that operate outside of low Earth orbit, the FLARE mission would further the readiness of these missions. This would aid in the reduction of future deep space CubeSat mission costs in addition to open up further possibilities with respect to deep space science and observation mission.

With respect to the environmental issues that arise due to space mission, the FLARE mission has an extremely low impact. Because this mission plans to utilize ride-sharing, it would not produce any extra pollution with regards to propellant emissions in the atmosphere. If the FLARE CubeSats are disposed by burning up into the atmosphere, they would produce an insignificant impact that would not likely be noticed at all. In the event of a launch failure, the spacecraft components are no more dangerous than those already included in the launch vehicle itself.

4.7 Ethical, Social and Health/Safety Issues

This mission is ethically and socially pertinent to improving propagation of near-Earth bodies. Due to the possible science return of this mission, the trajectories of near-Earth bodies such as asteroids could potentially be more accurately propagated. This could allow for early warning of meteorite impact areas and could play into future asteroid diversion missions by improving propagation estimates such that action could be taken earlier while small perturbations could significantly impact the asteroid’s course.

This mission would not likely impact the health and safety of the public. In spite of the FLARE CubeSats use of hydrazine as a propellant, the high technology readiness level (TRL) of the hydrazine systems mitigates the possible health impact to workers during this mission.

4.8 Manufacturability, Political and Global Impact Issues

Because this mission has been developed using standard, high TRL bus and components, it is easily manufacturable. Additionally, because the flyby altitudes are well over those considered to be in low Earth orbit, the probability of collision with other satellites is extremely small. This mission, however, does require the international cooperation of ground stations and possibly launch facilities. Of course, this also encourages and allows practice of such cooperation.

5.0 Design Critique

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

5.1 Strengths

The mission is well designed to constraints and goals. The mission design meets the requirements and should produce the data needed to evaluate the anomaly. The constraints of the mission are quite tight, especially the constraints of cost and trajectory. However, even with these constraints, solutions were derived.

High redundancy in data collection is also a strong point of the mission design. Telemetry data will include GPS measurements, ground station doppler/phase carrier determination observations, and satellite laser ranging. The conjunction of data will allow post-processing to filter the data for noise, bias, and perturbations.

5.2 Weaknesses

One design weakness is the requirement of ConOps A to launch into a Molniya type parking orbit. This may not be available as a ride-sharing opportunity, and a ROKOT launch may not be compatible with the SHERPA. As such, significant changes may need to be made to ConOps A should the requisite opportunities not be present.

The current design is preliminary and requires significant refinement if it is to be seriously considered for flight. Many simplifying assumptions are used in the preliminary analysis.

A significant weakness of the mission which bears further investigation is whether the flyby anomaly can be adequately solved by the application of higher order gravity models or the anisotropy of the speed of light to the current data set. If this is the case, the need outlined in the

mission scope may be largely erased, demotivating the mission. If this is the case, then this report may serve to increase motivation to investigate these explanations further, or to increase awareness of the resolution, which is not yet widespread.

5.3 Confidence

The confidence of the realizability of FLARE is relatively high. The FLARE report will serve as a baseline for any potential mission regarding testing of the flyby anomaly. Although the details of our baseline design are subject to change given further mission/system and alternatives development, the structure of the report and information contained will help in that refining process.

Areas of high confidence of success are the component selection, concept of operations, and mass/power/DV budgets. Trajectory confidence is moderate, since the approach is limited to patched conics and, as mentioned previously, patched conics may prove to be a poorer approximation than usual for this mission concept. Low confidence is placed in the radiation and link budgets due to lack of specific information used in their derivation.

5.4 Alternatives

Included in the body of this report are both mission and system design alternatives. The mission design alternatives include three preliminary ConOps (section 3.1) along with Primary and Secondary ConOps (section 3.1.5). The system design alternatives take the form of propulsion system alternatives (section 3.5.5) and data acquisition system alternatives (section 3.5.1), also given the implementation of a ConOps other than the Primary ConOps, further alteration from the baseline system designs (section 3.2 and 4.2) would be needed. One alternative, mentioned in section 5.2 above, is a GTO parking orbit applied to our baseline trajectory. This method's main advantage is that it there are many more launches to GTO than a Molniya type (highly inclined and eccentric) parking orbit with more ridesharing opportunities. If a GTO parking orbit is required, then this may suggest a re-evaluation of ConOps B

5.5 Remaining Design Refinements

This section will outline certain items initiated during the course of the design process which would be of immediate benefit as first steps should the mission be considered for further development.

5.5.1 CAD Model for Analysis

A CAD model was made using the battery, flight computer, EPS, power distribution system, and structure shown in section 4.2.2., in addition to the SGR-05U - Space GPS Receiver by Surrey Satellite Technology US LLC. and the VHF downlink / UHF uplink Full Duplex Transceiver by Innovative Solutions In Space. In order to assess the viability of a six unit cubesat

with components similar to those in section 4.2.2., this early CAD model, as seen in Figure 36, was developed.

Figure 36: 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.

5.5.2 Trajectory Refinement

Both baseline trajectories could benefit from re-development. The baseline for ConOps A uses patched conics approximations only. Preliminary attempts to use the outcome as an initial guess for a numerical integration approach emphasized the difficulty of the trajectory constraints. More accurate DVs that take into account perturbations and the errors associated

with patched conics, along with more accurate estimates of the magnitude of the trajectory correction maneuvers are needed.

ConOps B could be optimized by modifying the moon flyby, both in the B-Plane that is targeted, and by adding a burn to the flyby to improve the resulting geocentric orbit. This may well result in a higher magnitude of the anomaly expected, and reduce the DV requirements to within the capabilities of the SHERPA.

Finally, the exact timing of the trajectory could be improved to better match tracking needs and capabilities, alongside DSN/Estrack/TDRSS scheduling negotiation.

5.5.3 Comms Link Budget

A comms link budget for our system needs to be finalized for the various scenarios during which communication is needed. These scenarios include gathering navigation/trajectory data during the flyby phases from Earth’s sphere of influence (0.0062 AU or ~1 million km) through closest approach (~1500 km), as well as communicating as an uplink only, trajectory correction maneuvers during the heliocentric trajectory phases (~0.01 AU). A preliminary comms link budget is outlined in section 4.2.5.

6.0 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 formulas that approximate the anomaly magnitudes have been contested by two dominant explanations for the anomaly, high order gravity effects and the anisotropy of the speed of light. The data collected by FLARE will further the investigation of describing the anomaly by observation of trajectory details, as well as the two aforementioned 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 (~0.74) 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 and 70% exposure for deployed 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, 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 Earth bound systems cooperating with Earth orbiting satellites (e.g. GPS) and the satellite experiencing the anomaly. The Earth based systems include the Global Positioning System (GPS) and radio doppler monitoring via DSN stations and/or Estrack stations, and possibly Satellite Laser Ranging capable ground stations as a compliment or substitute for GPS. The Earth orbiting satellites include TDRSS (Tracking and Data Relay Satellite System) and GNSS (Global Navigation Satellite System). The satellites experiencing the anomaly are the FLARE CubeSats and potentially the exit assist vehicle (SHERPA 2200). The FLARE CubeSats subsystems that gather trajectory, e.g. identify the anomaly, are the Communications, GPS, and SLR subsystems. The components chosen to satisfy those roles are an Iris X-Band Transponder, a FOTON GPS receiver and a passive SLR reflector. A hydrazine motor capable of ~150 m/s was chosen to suit trajectory correction needs. The projected cost for this mission, excluding the cost associated with the exit assist vehicle (SHERPA 2200), is somewhat greater than $4.5 million.

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. High order gravity terms (modeling up to 100X100) 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. Applying these gravity models and the anisotropy of the speed of light models to all the heritage missions which experienced the anomaly could serve to further evaluate their validity.

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.

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.[13] Aerojet Rocketdyne. “CubeSat Modular Propulsion Systems (MPS)”,https://www.rocket.com/cubesat.[14] Surrey Satellite Technology US LLC. “SGR-05U – Space GPS Receiver”, http://www.sst-us.com/shop/satellite-subsystems/gps/sgr-05u- space-gps-receiver.[15] Bill Schreiner, Doug Hunt, Chris Rocken, Sergey Sokolovskiy. “Approach and Quality Assessment of Precise GPS Data Processing at the UCAR CDAAC”, University Corporation for Atmospheric Research (UCAR)COSMIC Project OfficeBoulder, CO[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.[22] JPL “Basics of Space Flight” Section II Chapter 13 Spacecraft Navigation. http://www2.jpl.nasa.gov/basics/bsf13-1.php[23] Srinivisan, Dipak K., and Fielhauer, Karl B., “The Radio Frequency Subsystem and Radio Science on the MESSENGER Mission”, August 2007. http://www-geodyn.mit.edu/srinivasan.mercuryrs.ssr07.pdf

[24] Taylor Jim, et al., “Galileo Telecommunications”, DECANSO Design and Performance Summary Series, Article 5, JPL, July 2002. http://descanso.jpl.nasa.gov/DPSummary/Descanso5--Galileo_new.pdf[25] Spaceflight, Inc. Secondary Payload Users Guide. 3415 S. 116th St, Suite 123Tukwila, WA 98168. SF-2100-PUG-00001, Rev D 2013-03-05.[26] Mukai, Ryan et al., "Juno Telecommunications", DECANSO Design and Performance Summary Series Article 16, JPL, October 2012.[27] 2002367B Payload Spec for 3U 6U 12U 27U. Planetary Systems Corporation, 21 July, 2014.[28] Adler, Stephen L. “Modeling the Flyby Anomalies with Dark Matter Scattering.” Princeton Institute for Advance Study, 17 Feb. 2012. Web. <http://arxiv.org/pdf/1112.5426.pdf> [29] Robertson, R., Shoemaker, Michael. “Highly Physical Penumbra Solar Radiation Pressure Modeling and the Earth Flyby Anomaly”. SpaceOps Conferences, 5-9 May 2014. Web. <http://arc.aiaa.org/doi/pdf/10.2514/6.2014-1881>. [30] McCulloch, M.E. “Can the Flyby Anomalies Be Explained by a Modification of Inertia?”. Journal of British Interplanetary Society, 18 Dec. 2007. Web. <http://arxiv.org/pdf/0712.3022v1.pdf> [31] Mbelek, Jean P. “Special Relativity May Account for the Spacecraft Flyby Anomalies.” Service D’Astrophysique, 15 Mar. 2009. Web. <http://arxiv.org/ftparxiv/papers/0809/0809.1888.pdf>[32] Atchison et al. “Lorentz Accelerations in the Earth Flyby Anomaly”. Journal of Guidance, Control, and Dynamics. 2012. Web. <http://arc.aiaa.org/doi/pdf/10.2514/1.47413>[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.[34] Duncan, Courtney, “Microwaves: Communications and Navigation in Deep Space … even in nano-SpaceCraft”. San Bernardino Microwave Society. Corona, California. Oct 2, 2014. [35] Courtney Duncan and Amy Smith, “Iris Deep Space CubeSat Transponder”. Jet Propulsion Laboratory, California Institute of Technology. CubeSat Workshop #11, Cal Poly San Luis Obispo. April 23, 2014.[36] Thompson et al., “Reconstruction of Earth Flyby by the JUNO Spacecraft”. California Institute of Technology, 2014. Web.[37] NovAtel. “OEM628 Triple-Frequency + L-Band GNSS Receiver”,http://www.novatel.com/prodecuts/gnss-receivers/oem-receiver-boards/oem6-receivers.[38] European Space Agency. “SAC-C (Satelite de Aplicaciones Cientificas-C)”,https://directory.eoportal.org/web/eoportal/satellite-missions/s/sac-c.[39] Orfeu Bertolami, Frederico Francisco, Paulo J. S. Gil, Jorge Paramos. “Testing the Flyby Anomaly with the GNSS Constellation”. WSPC/Instruction file, arSiv:1201.0163v1 [physics.space-ph]. Universidade T´ecnica deLisboa. Lisboa, Portugal. Jan 4, 2012.[40] General Dynamics. “Small Deep-Space Transponder (SDST)”. http://www.gd-ais.com/Documents/Space%20Electronics/SDST%20-%20DS5-813-12.pdf[41] Tyvak. Intrepid System Board. “http://tyvak.com/intrepidsystemboard/”[42] Antenna Development Corporation. “Microstrip patch Antennas”. http://www.antdevco.com/ADC-0509251107%20R6%20Patch%20data%20sheet_non-ITAR.pdf[43] Sara Spangelo, Matthew Bennett, Daneil Meinzer, Andrew Klesh, Jessica Arlas, James Cutler. “Design and Implementation of the GPS Subsystem for the Radio Aurora Explorer”. University of Michigan, 1320 Beal Ave, Ann Arbor, MI 48109. Jan. 7, 2013.[44] Cahill, R.T. “Resolving Spacecraft Earth-Flyby Anomalies with Measured Light Speed Anisotropy”. School of Chemistry, Physics and Earth Sciences, Flinders University, Adelaide 5001, Australia. July, 2008.

[45] Duncan, Courtney. “Iris for INSPIRE CubeSat Compatible, DSN Compatible Transponder. Flight Communications Systems Section 337. Jet Propulsion Laboratory, California Institute of Technology. July 31, 2013.[46] Adler, Stephen L. “Can the Flyby Anomaly be Attributed to Earth-Bound Dark Matter?”. Institute for Advanced Study, Einstein Drive, Princeton, NJ 08540, USA. Dec, 2008.[47] Busack, H.J. “Expected Velocity Anomaly for the Earth Flyby of Juno Spacecraft on October 9, 2013”. Wulfsdorfer Weg 89, 23560 Lübeck, Germany. September 25, 2013 [48] Courtney B. Duncan, Amy E Smith and Fernando H. Aguirre. “Iris Transponder - Communications and Navigation for Deep Space”, SSC14-IX-3. Jet Propulsion Laboratory, California Institute of Technology. 28th Annual AIAA/USU Conference on Small Satellites.[49] Joplin, A.J., E.G. Lightsey, T.E. Humphreys, "Development and Testing of a Minaturized, Dual-Frequency GPS Receiver for Space Applications," Proceedings of ION ITM, Newport Beach, California, 2012.[50] Pumpkin Incorporated, “Pumpkin Price List,” http://www.pumpkininc.com/content/doc/forms/pricelist.pdf.[51] Delft University of Technology, “Propulsion system cost data,” http://www.lr.tudelft.nl/en/organisation/departments/space-engineering/space-systems-engineering/expertise-areas/space-propulsion/design-of-elements/cost/.[52] Brumbaugh, K.M., “The Metrics of Spacecraft Design Reusability and Cost Analysis as Applied to CubeSats,” Thesis Presented to the Faculty of the Graduate School of The University of Texas at Austin, The University of Texas at Austin, May 2012.[53] Andrew Klech, John Baker, Julie Castillo-Rogez, Lauren Halatek, Neil Murphy, Carol Raymond, Brent Sherwood, John Bellardo, James Cutler and Glenn Lightsey. “INSPIRE: Interplanetary NanoSpacecraft Pathfinder In Relevant Environment”, SSC13-XI-8. 27th Annual AIAA/USU Conference on Small Satellites.[54] Andrew Kwas, Eric MacDonald, Dan Muse, Ryan Wicker, Craig Kief, Jim Aarestad, Mike Zemba, Bill Marshall, Carol Tolbert and Brett Connor. “Enabling Technologies for Entrepreneurial Opportunities in 3D Printing of SmallSats”, SSC14-III-7. 28th Annual AIAA/USU Conference on Small Satellites.[55] spaceflightservices.com. “Pricing”. <http://spaceflightservices.com/pricing-plans/>.[56] SLR. “Satellite Laser Ranging”. NASA. May 4, 2015, <http://esc.gsfc.nasa.gov/space -communications/NEN/slr.html>.[57] “Satellite Laser Ranging and Earth Science”. NASA Space Geodesy Program. May 4, 2015. <http://ilrs.gsfc.nasa.gov/docs/slrover.pdf>.

7.1 Image References<https://thelistlove.files.wordpress.com/2014/03/26.jpg><http://darkroom.baltimoresun.com/wp-content/uploads/2012/05/AFP_Getty-TOPSHOTS-US-SPACE-INDUSTRY-FALCON-9.jpg><http://spaceflightservices.com/wp-content/uploads/2013/08/SHERPA_w_panels_v002.png><http://www.nasa.gov/sites/default/files/thumbnails/image/dellingr_artist_concept.jpg><http://i.space.com/images/i/000/025/089/i02/orion-service-module-engine-burn.jpg?1358369866><http://www.esa.int/var/esa/storage/images/esa_multimedia/images/2003/07/binary_system_earth-moon/102 2 5612- 2-eng- GB/Binary_system_Earth-Moon.jpg><http://i.ytimg.com/vi/FjCKwkJfg6Y/maxresdefault.jpg><https://icubesat.files.wordpress.com/2014/06/icubesat-org_2014_b-1-4-kupda_sauder_20140617.pdf><http://inspirehep.net/record/833373/plots>< http://esc.gsfc.nasa.gov/assets/images/TLRS-4%205-09.jpg>

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

Figure: SHERPA configuration [3, 25].

Figure: SHERPA propulsion capabilities [3].

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

Figure: RAX CubeSat, GPS heritage section 3.4.4 and Figure 10, GPS link budget [43].

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

Figure: CSD payload specifications sheet [27].

Figure: CSD specifications [4].

Anomaly Visualization: MATLAB plots

Doppler Noise Magnitudes courtesy of JPL [20]

INSPIRE configuration using the JPL developed LMRST [45]

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 StrengthsThe 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 KangAmritpreet 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 AlfaroJeffrey 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 KangThe 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 AlfaroThe 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, have performed 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 RamseyThese 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 is 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

Figure 12: FLARE Initial Personnel Assignments

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 Statements7.1 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 AlfaroIn 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.

I performed the trade studies for trajectory, tracking, and separation, as well as prepared the baseline trajectories for both ConOps, as well as the comparison between the ConOps. I also worked on the design critique and risk analysis as well as the finalized version of the mission scope.

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 RamseyAt the initiation of the mission design process it made sense for me, as systems engineer,

to manage and 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 initial 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 summarized 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 and system vs. low level system requirements. The new sections that I wrote include: Day in the life of FLARE, Requirements Traceability Matrix, part of JUNO Findings, Prospective Modeling Analysis, Comms Link Budget and EbNo Analysis, and most of the Design Critique section.

7.4 Kyle ChaffinMy 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. I performed a volume analysis of the spacecraft components in order to ensure that components will fit in 6U CubeSat.

7.5 Anthony HuetWith my writing skills and general knowledge of spaceflight I wrote the introduction and

updated and edited the final version of it. I carried out the propulsion subsystem trade study, comparing the various propulsion methods available to the CubeSat and choosing the best. The study considered monopropellant, bipropellant, electric and cold gas thruster. A calculation of delta-V from electric propulsion was also performed. I determined the power requirements for each component and created the power equipment list (PEL) to provide an overview of the power usage of the spacecraft. Three different scenarios were also created to analyze various situations the spacecraft may be in and how power usage would look in those.

I also researched the Satellite Laser Ranging system (SLR) to provide information on its capabilities and using it for tracking the CubeSat. I analyzed the risks present for this mission and created the risk register and risk ratings that evaluate the probability, severity, and possible mitigations for various risks.

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 CubeSat’s’ 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.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 after design refinements.

Appendix IV: JPL feedbackJPL sticky note feedback note: this section of feedback was handwritten and transcribedProgrammaticsWhy 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 BCan 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 perturbations

Juno: 10X10>>50X50 6-7mm/s dV @ periapsis.

Travis Imken FeedbackFLARE A-Team SessionMarch 16, 2015 9a – 12:15p

JPL 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 hydrazineNeed 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 thereBalloonsLasersGet 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