Conference PaperPDF Available

A Heliocentric Satellite Constellation for Continuous Solar Coverage and Space Weather Monitoring

March 2024

DOI:10.1109/AERO58975.2024.10521338

Conference: 2024 IEEE Aerospace Conference
At: Big Sky, Montana, USA

Authors:

Allan Shtofenmakher

Massachusetts Institute of Technology

Daniel Gochenaur

Massachusetts Institute of Technology

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While the Sun provides the Earth with the energy needed to sustain life, the volatility associated with this intense energy source generates solar weather, which can have devastating implications on Earth. Solar weather can result in data compromise, radio interference, premature satellite deorbit, and even failure of the power grid. To mitigate the negative effects of solar weather, constant observation of the entirety of the Sun’s surface is essential. This complete picture of the Sun’s ever-changing state will help scientists anticipate solar events that may negatively impact life on Earth. A heliocentric satellite constellation called the Solar Unobstructed Network-based First Long-term Outer-space Weather Effects Research (SUNFLOWER) Observatory is proposed to continuously monitor coronal mass ejections, sunspots, and coronal holes with a suite of science instruments capable of collecting data in various electromagnetic wavelengths. This report offers a holistic view of the mission and spacecraft architectures. The paper begins with a discussion of motivation, mission objectives, and influential past missions. Next, a high-level overview of the mission design flow, mission-level requirements, and cost and schedule estimation assumptions is explored. This is followed by an analysis of the stakeholders and associated value flows and identification of system boundaries. Next, high-level design decisions for critical components of the system architecture and project risks and risk mitigation strategies are discussed. Results for instrument selection, constellation design, and spacecraft design are presented along with the reasoning behind the recommended architectures and design decisions. The final result is an estimate of the overall mission cost and schedule—roughly $4B in FY2025 USD over an 18-year lifecycle beginning in FY2025. The conclusion summarizes the proposed constellation, composed of nine identical spacecraft—each containing a magnetograph, an extreme ultraviolet imager, and a coronagraph—in a Walker-Delta 54.7° configuration at one AU, with three spacecraft in each of three planes. This solution offers continuous 4π-steradian remote sensing coverage of the solar surface—including the poles—with daily communication of science and state-of-health data over Ka-band frequencies to Earth using 34-m ground stations within the Deep Space Network (DSN). To circumvent the significant burden that would be placed on DSN, a compelling and mutually beneficial case for investing in additional 34-m antennas is presented. The paper concludes with recommendations for future work on the SUNFLOWER Observatory.

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A Heliocentric Satellite Constellation for Continuous

Solar Coverage and Space Weather Monitoring

Allan Shtofenmakher

ashtofen@mit.edu

Duncan Miller

dmiller1@mit.edu

Katelyn Sweeney

ksweeney@g.harvard.edu

Frederick Ajisafe

ajisafe@mit.edu

Leilani Trautman

trautman@mit.edu

Daniel Gochenaur

dgochena@mit.edu

Luke de Castro

lukedc@mit.edu

Joel Jurado Diaz

jurado@mit.edu

Akila Saravanan

akilasar@mit.edu

Nadia Khan

nadiak@mit.edu

Department of Aeronautics and Astronautics

Massachusetts Institute of Technology

Cambridge, MA

Benjamin Waters

bwatr529@mit.edu

Tai Zheng

taizheng@alum.mit.edu

Alexis Lepe

alepe@mit.edu

Joana Nikolova

jonik@mit.edu

Olivier L. de Weck

deweck@mit.edu

Robert Cato III

rcato3rd@alum.mit.edu

Alexander Koenig

koe@alum.mit.edu

Claire McLellan-Cassivi

cmclcas@mit.edu

Clara Ziran Ma

czm6@mit.edu

Edward F. Crawley

crawley@mit.edu

Abstract—While the Sun provides the Earth with the energy

needed to sustain life, the volatility associated with this intense

energy source generates solar weather, which can have devas-

tating implications on Earth. Solar weather can result in data

compromise, radio interference, premature satellite deorbit, and

even failure of the power grid. To mitigate the negative effects

of solar weather, constant observation of the entirety of the

Sun’s surface is essential. This complete picture of the Sun’s

ever-changing state will help scientists anticipate solar events

that may negatively impact life on Earth. A heliocentric satel-

lite constellation called the Solar Unobstructed Network-based

First Long-term Outer-space Weather Effects Research (SUN-

FLOWER) Observatory is proposed to continuously monitor

coronal mass ejections, sunspots, and coronal holes with a suite

of science instruments capable of collecting data in various elec-

tromagnetic wavelengths. This report offers a holistic view of the

mission and spacecraft architectures. The paper begins with a

discussion of motivation, mission objectives, and inﬂuential past

missions. Next, a high-level overview of the mission design ﬂow,

mission-level requirements, and cost and schedule estimation

assumptions is explored. This is followed by an analysis of

the stakeholders and associated value ﬂows and identiﬁcation of

system boundaries. Next, high-level design decisions for critical

components of the system architecture and project risks and risk

mitigation strategies are discussed. Results for instrument selec-

tion, constellation design, and spacecraft design are presented

along with the reasoning behind the recommended architectures

and design decisions. The ﬁnal result is an estimate of the

overall mission cost and schedule—roughly $4B in FY2025 USD

over an 18-year lifecycle beginning in FY2025. The conclusion

summarizes the proposed constellation, composed of nine iden-

tical spacecraft—each containing a magnetograph, an extreme

ultraviolet imager, and a coronagraph—in a Walker-Delta 54.7°

conﬁguration at one AU, with three spacecraft in each of three

planes. This solution offers continuous 4π-steradian remote

sensing coverage of the solar surface—including the poles—with

daily communication of science and state-of-health data over

Ka-band frequencies to Earth using 34-m ground stations within

the Deep Space Network (DSN). To circumvent the signiﬁcant

burden that would be placed on DSN, a compelling and mutually

beneﬁcial case for investing in additional 34-m antennas is pre-

sented. The paper concludes with recommendations for future

work on the SUNFLOWER Observatory.

TABLE OF CONTENTS

I. INTRODUCTION ................................... 1

II. METHODOLOGY ................................. 2

III. ANALYSI S ....................................... 3

IV. RES ULTS ......................................... 5

V. CONCLUSION ..................................... 11

APPENDICES......................................... 12

A. REQUIREMENTS MATRI X ........................ 12

B. RISK MATRIX .................................... 12

C. SCIENCE TRAC EABIL IT Y MATR IX ............... 12

D. GANTT CH ART ................................... 12

ACKNOWLEDGMENTS ............................... 12

REFERENCES ........................................ 12

BIOGRAPHY ......................................... 15

I. INTRODUCTION

A. Motivation

The intense energy radiated by the Sun is a constantly

evolving phenomenon which induces ever-changing mag-

netic ﬁelds. These changes cause variances in and expulsions

of plasma and radiation that have serious implications for

life on Earth. Solar radiation ﬂares can cause radio blackout

storms that degrade radio communication at high frequencies.

Coronal mass ejections (CMEs), characterized by large ejec-

tions of plasma from the Sun’s corona, induce geomagnetic

storms in Earth’s magnetic ﬁeld. When these CMEs reach

the Earth’s surface, they can induce surplus current in power

grids and cause blackouts [1], [2]. For example, a geo-

magnetic storm in 2003 caused an hour-long blackout across

Scandinavia, as well as the rerouting of many aircraft around

the world [3]. Geomagnetic storms can also impact radio

signals and degrade GPS performance [1], [2]. These are only

a few of the ways the Sun can negatively affect life on—and

the environment around—Earth. The National Oceanic and

Atmospheric Administration (NOAA) monitors the Sun and

predicts space weather events by observing solar ﬂares and

CMEs. This is accomplished via a host of sensors on the

Earth’s surface and in space [4]. While these observations

help to predict and react to solar weather, a lack of data

regarding solar and magnetic variability at the Sun’s poles and

on the far side of the (rotating) Sun curtails heliophysicists’

understanding of this celestial body. For example, the solar

magnetic ﬁeld, which shapes and inﬂuences phenomena like

CMEs, originates from the Sun’s poles [5]. Likewise, views

of the far side of the Sun allow improved tracking of solar

weather indicators like sunspots [6].

B. Objectives

Continuous 4π-steradian (sr) coverage of the Sun’s surface

through a heliocentric constellation with polar and far-side

observations will improve space weather predictions and

overall understanding of the Sun. The Solar Unobstructed

Network-based First Long-term Outer-space Weather Effects

Research (SUNFLOWER) Observatory seeks to improve un-

derstanding of the magnetic variability of the Sun as it relates

to space weather events and understanding of other stars by

continuous observation of the Sun’s entire surface, including

the poles, using a constellation of heliocentric spacecraft. The

overarching scientiﬁc objectives of this system are: 1) support

global heliophysics research and 2) provide more accurate

space weather forecasting and monitoring.

The ﬁrst objective involves providing 4πsr of coverage of the

solar surface, to include the solar poles, generating continu-

ous high-cadence data of the entire solar surface, and provid-

ing a better understanding of the Sun’s magnetic ﬁeld. The

second objective includes monitoring CMEs and sunspots,

collecting high-resolution magnetic data, and producing an

overall increased data return rate from past missions.

C. Past Missions

SUNFLOWER is unprecedented in its size as a heliocentric

constellation. However, multiple standalone missions provide

inspiration and valuable lessons to the proposed constellation.

Between 1994 and 1995, the Ulysses spacecraft, developed

by the European Space Agency (ESA) in conjunction with

the National Aeronautics and Space Administration (NASA),

became the ﬁrst spacecraft to observe the Sun’s poles. The

Ulysses spacecraft used a gravity assist around Jupiter to

maneuver out of the ecliptic plane and reach over 80° latitude

relative to the Sun. These polar passes only spanned a

few months each, with years of relative idleness between

[7]. SUNFLOWER improves upon Ulysses by providing

continuous polar observations.

A more recent mission to pursue polar observations of the

Sun is the ESA Solar Orbiter, which launched in 2020 and

will eventually reach 33° inclination above the ecliptic plane

[8], [9]. This spacecraft uses multiple gravity assists from

Venus to reach its desired inclination over the course of

several years [10]. Like Ulysses, with only one spacecraft,

these observations are not continuous. Additionally, Solar

Orbiter spends a very limited amount of time in view of the

Sun’s highest latitudes. SUNFLOWER improves upon this

by observing all latitudes of the Sun continuously.

Another inspirational mission is NASA’s Solar Dynamics

Observatory (SDO). The launch of SDO represented another

large step toward understanding the structure of the Sun’s

magnetic ﬁeld [11]. However, since its launch in 2010, its

instruments continue to undergo severe degradation. For ex-

ample, the Extreme Ultraviolet Variability Experiment (EVE)

has already experienced charge-coupled device degradation

of 90% at the 105 nm wavelength [12]. This loss of per-

formance contributes to the motivation for a constellation of

spacecraft with newer instruments that degrade at slower rates

than past missions, which is discussed in detail in Section IV.

Combining these points of improvement and novelty, SUN-

FLOWER seeks to improve upon past missions by contin-

uously covering the poles—in contrast to time-limited ﬂy-

bys—which is accomplished using a multi-plane heliocentric

constellation with 4π-sr coverage. It also seeks to account

for instrument degradation in missions like SDO by using

newer instruments that are inherently resilient to degradation.

In addition, SUNFLOWER will utilize alternative trajectory

methods to interplanetary ﬂybys—namely, advanced solar

electric propulsion (SEP, discussed further in Section IV).

The proposed concept of operations (CONOPS) for the SUN-

FLOWER Observatory, beginning with initial launch to low

Earth orbit (LEO), is presented at a high level in Figure 1 and

is also discussed further in Section IV.

Figure 1.SUNFLOWER CONOPS.

II. METHODOLOGY

The respective methodologies for the systems design process,

requirements derivation, and costing and scheduling process

on this program follow.

A. Systems Process

To maximize the value of SUNFLOWER, the Systems En-

gineering (SE) team began the design of this mission with

a detailed stakeholder analysis. Once the stakeholders were

identiﬁed, key science objectives were determined to meet

their needs. From this, a list of solar properties of interest,

ideal spacecraft distance from the Sun and necessary tech-

nology payloads to collect the science data of interest were

identiﬁed. By comparing metrics such as complexity, cost,

solar coverage, and science value, a constellation architecture

was down-selected and used to inform spacecraft trajectories,

core constellation design, and types of propulsion. Given

these speciﬁcations, a spacecraft bus design could be created

with thermal, power, communications, attitude determination

and control system (ADCS), material, and structural trade

studies. Finally, given the spacecraft mass, the ideal launch

vehicle was selected from a trade study including vehicle

cost, risk, and capability. Figure 2 shows the methodology

followed by the team.

Figure 2.Systems process ﬂowchart.

B. Requirements Derivation

The process for deriving high-level requirements for the

SUNFLOWER program was largely driven by inputs pro-

vided by key stakeholders. In particular, many are derived

from the overarching SUNFLOWER description and mission

objectives, such as 4π-sr coverage of the Sun, space weather

monitoring, and continuous communication of science data

back to Earth. As an example, the requirement for mission

lifetime was deﬁned to align with the average duration of a

solar cycle—roughly eleven years [13]. Other factors affect-

ing the derivation of the requirements include cost, schedule,

and practical considerations—such as compatibility with the

NASA Deep Space Network (DSN)—to enable the mission

as a whole.

Table 15 in Appendix A presents the resulting requirements

matrix, which includes eleven requirements in all.

C. Costing and Scheduling Process

The overall program schedule, with the exception of instru-

ment development, is primarily derived from the schedules

of similar missions, such as Ulysses and Parker Solar Probe

[14], [15]. Each payload instrument selected for use in the

SUNFLOWER program has previously ﬂown [16], is set to

ﬂy in the near future [17], or is space qualiﬁed and ready

to launch [18]. This extensive heritage enables an accel-

erated timeline for instrument development relative to other

heliophysics programs. This, in turn, enables the operating

assumption that the payload development timeline does not

drive the overall program schedule and can largely be en-

closed within the timeframe for spacecraft/bus development.

The preliminary cost estimate for the SUNFLOWER program

was derived using information from several sources [14],

[19], [20], [21], [22], [23], [24].

In particular, cost estimating relationships (CERs) from [19],

often called “SMAD” or “SME,” were used to translate key

metrics—such as subsystem-level dry mass estimates, sci-

ence instrument power draws, and mission lifetime——into

cost estimates for the program’s work breakdown structure

(WBS). The uncertainty surrounding these input parameters

was used to derive low, medium, and high cost estimates for

each WBS line item. Since the cost models in [19] assume

FY2010 USD, the results were adjusted for inﬂation.

The high-level costing models in SME combine launch and

orbital operation costs into one estimate [19]. However, for

SUNFLOWER, the Constellation Design and Astrodynamics

(CDA) team generated reasonable launch cost estimates as

part of their launch vehicle trade study, and these ﬁgures were

selected for the costing model due to their higher ﬁdelity [20],

[21]. To complete the set, mission operation and data analysis

costs were determined using NASA’s Mission Operations

Costing Tool (MOCET) for a large Near-Earth Discovery

Heliophysics program operating under the SUNFLOWER

timeline [22].

A qualiﬁcation or “qual” unit is also included in the cost

estimate. Although not strictly necessary, a qual unit enables

engineers to verify that the spacecraft will survive the launch

vehicle and space environments while there is still time to

adjust the ﬁnal spacecraft design. For a program that is

expected to ﬂy multiple spacecraft far beyond Earth, the

reduction in risk associated with the qualiﬁcation unit is likely

well worth the added cost. Further discussion of cost and

schedule is presented in Section IV.

III. ANALYSIS

Analyses are presented for the stakeholder investigation, the

system boundary deﬁnition, and a high-level overview of

instrument selection, constellation design, satellite bus archi-

tecture, and risks.

A. Stakeholder Investigation and Value Network

As a NASA project, this system is at the center of a com-

plex stakeholder network that includes the U.S. government,

NOAA, and the heliophysics community. Figure 3 shows the

Stakeholder Value Network created from the SUNFLOWER

observatory and its stakeholders. The most important value

ﬂows—in red—are critical to the program’s survival. Yellow

ﬂows reﬂect values affecting the project performance, and

green ﬂows reﬂect excitements that are desirable but not

strictly necessary for project success.

As a civil space agency, NASA’s relationship and value

exchanges with the government (i.e., the Executive Branch

and Congress) are particularly important for the project’s suc-

cess. Speciﬁcally, Congress is the source of program policy,

approval, and funding, all of which are vital for the system

to exist. Other necessary value ﬂows include materials that

come from manufacturers in the commercial space industry

and scientiﬁc results that are provided by the heliophysics

program scientists using SUNFLOWER Observatory data.

Prioritizing stakeholders by value ﬂows indicates that the

U.S. government, commercial space industry, and the helio-

physics community—in particular, the NASA Heliophysics

Science Mission Directorate—are the most important for

the SUNFLOWER program. The heliophysics community

is considered essential for having shaped SUNFLOWER’s

scientiﬁc objectives. Speciﬁcally, discussions with stakehold-

ers revealed that the core science needs driving the goals

of this project are novel solar observation points outside of

the Earth-Sun line and a better understanding of the solar

magnetic ﬁeld. Indeed, these two needs are connected,

as continuous full-surface observations—especially of the

Sun’s poles—are required to determine the magnetic ﬂux

boundaries and the magnetosphere’s role in shaping solar

weather. Accordingly, SUNFLOWER will seek to offer the

science community never-before-collected continuous full-

surface magnetograms, extreme ultraviolet (EUV) imaging,

and corona imaging, which will provide critical insights into

the behavior and boundaries of the Sun’s magnetic ﬁeld. Full-

surface magnetograms inclusive of the poles will allow the

science community to determine the exact ﬂux boundaries of

the Sun’s magnetic ﬁeld, which will substantially increase he-

liophysicists’ understanding of the creation of solar weather,

as well as the behavior of the various layers of the Sun. EUV

and coronal imaging will complement these magnetograms

by providing detailed images of solar ﬂares and sunspots.

SUNFLOWER’s emphasis on space weather reveals notable

value ﬂows among itself, NOAA, the energy industry, and the

U.S. government. A value loop is created when observations

from SUNFLOWER go to NOAA, which uses them to make

space weather predictions for the energy industry. The energy

industry uses these predictions, along with the government’s

space weather mitigation policy, to provide reliable electricity

to the general public. In turn, the public provides addi-

tional support for the SUNFLOWER program to Congress,

encouraging potential program extension after the primary

mission. This connection with the energy industry is a

unique characteristic that sets the SUNFLOWER stakeholder

landscape apart from that of most other space programs.

Figure 3.Stakeholder Value Network.

B. System Boundaries

Figure 4 deﬁnes the elements of the SUNFLOWER system

and its context in a largely solution-neutral manner. The over-

all system consists of an arbitrary number (N) of spacecraft

(SC) contained within an arbitrary number (Z) of constella-

tion planes, which gather energy and collect science data from

the Sun. Each SC interacts with its local space environment

by exchanging heat, encountering charged particles, enduring

radiation, and more. Likewise, each SC communicates with

ground stations on Earth to transmit science and state-of-

health data and to receive commands from operators.

Figure 4.System boundary deﬁnition for the

SUNFLOWER Observatory program.

C. High-Level Design: Instrument Selection

The preliminary scientiﬁc instrument package for SUN-

FLOWER was selected as a result of discussions between the

Heliophysics and Solar Monitoring (HSM) team and various

stakeholders that helped match overall scientiﬁc objectives to

possible scientiﬁc payloads.

Starting from the core scientiﬁc objectives, the following

goals were determined through stakeholder discussions: 1)

determine the solar magnetic ﬂux boundaries, and 2) de-

termine the primary drivers of solar wind. Conversations

were held with stakeholders in the heliophysics community to

determine the observable solar phenomena needed to address

scientiﬁc objectives. Observing CMEs, sunspots, and coronal

holes continuously over the entire surface would fulﬁll these

overarching goals. Speciﬁcally, the ﬁrst goal can be fulﬁlled

through observations of the magnetic ﬂux around coronal

holes (e.g., via omni-surface magnetograms), and the second

goal can be fulﬁlled by measuring the magnetic ﬂux, the

structure of the corona, and the locations of sunspots (e.g.,

via full-disk photospheric magnetograms, continuous EUV

imaging, and coronagraphic imaging).

Discussions with stakeholders determined the types of in-

struments that would be used to observe these phenomena.

Instrument selection is discussed in detail in Section IV.

D. High-Level Design: Constellation & Astrodynamics

Balancing heliophysics science goals, spacecraft quantity,

and bus propulsion needs drove the CDA team’s design pro-

cess for the constellation architecture. The architecture must

consider a feasible bus design given current launcher and

in-space propulsion technologies. The relevant trade spaces

include the quantity and placement of ﬁnal constellation

orbits along with their corresponding deployment trajectories.

The SUNFLOWER Observatory prioritizes remote sensing

over in-situ heliophysics measurements. Therefore, the ﬁ-

nal orbit trade space was narrowed to circular orbits at 1

astronomical unit (AU) to achieve continuous 4π-sr coverage

while minimizing cost and fuel mass. Adjusting either the

eccentricity or semimajor axis would raise ∆vand require

more stringent thermal and radiation considerations without

substantially improving coverage or science value.

With the orbit shape selected, a variety of orbital plane

inclinations and number of satellites per plane were analyzed.

Each proposed constellation design was evaluated for the

scientiﬁc value it would deliver, with speciﬁc focus on the

worst-case viewing angles offered of each point on the Sun’s

surface. Section IV provides a detailed discussion on the

selected orbital plane conﬁguration.

The technical cost of each ﬁnal orbit was determined from

the ∆vand time of ﬂight needed to ﬂy the full trajectory

from Earth. The trade study included chemical propulsion,

electric propulsion, and solar sails, all considered both with

and without gravity assists.

From investigations into historic heliophysics missions and

the degradation of their instruments over time, combined with

recent advancements in radiation shielding technology, the

reasonable expected lifetime of each spacecraft was estimated

to be approximately 11 years. To maximize the time each

spacecraft spent in its ﬁnal orbital conﬁguration, solar sails

were eliminated from the study due to their high time of ﬂight

and relatively low Technology Readiness Level (TRL).

Likewise, gravity assists also required multi-year times of

ﬂight. A trajectory utilizing a Jupiter gravity assist (JGA)

would permit a ballistic inclination increase with no addi-

tional ∆v. Similar mission concepts have explored a JGA,

including the Solar Polar Observing Constellation (SPOC)

[25]. The analysis in [25] agrees with the CDA team’s

JGA investigation that the total time of ﬂight would exceed

six years. Given the additional complexity of deep space

operations around Jupiter, the additional delay before useful

science operations, and the availability of a feasible direct

SEP injection, electric propulsion emerged as the selection

from this trade study. The JGA remains as a backup trajectory

in case of a substantial reduction to the baseline ∆vbudget.

After the Conceptual Design Review (CoDR), the CDA team

focused on higher-ﬁdelity modeling of the proposed Walker-

Delta 54.7◦constellation. This included investigations into

launch vehicle capabilities, low-thrust trajectory optimiza-

tion, and sensitivities to launch date and trajectory error. The

results of these analyses are discussed in Section IV.

Trajectory design using low-thrust propulsion does not lend

itself to optimal analytical solutions. The CDA team devel-

oped feasible hand-designed inclination and phasing maneu-

vers between 0.8 AU and 1.1 AU with the help of simulations

from NASA’s open-source General Mission Analysis Tool

(GMAT) [26]. The results from the GMAT analysis are

treated as baseline for all subsequent analyses. In addi-

tion, the team developed a preliminary nonlinearly optimized

trajectory with pykep, a Python package for low-thrust

optimization [27], which yielded improvements in both ∆v

and time of ﬂight. Further trajectory optimization is recom-

mended as future work.

E. High-Level Design: Spacecraft Bus

The goal of the SUNFLOWER bus team is to provide a

resilient architecture to support the scientiﬁc objectives of

the SUNFLOWER Observatory mission. Critical subsystems

include command and data handling (C&DH), attitude de-

termination and control, communications, electrical power,

thermal management, and structures and mechanisms. The

spacecraft bus architecture trade space—including radioiso-

tope thermoelectric generators (RTGs) for power—is shown

in Table 1. From these options, the spacecraft is designed to

satisfy constraints and requirements determined by the SE,

HSM, and CDA teams. These are shown in Table 2.

The communications system design depended on available

infrastructure and technologies (both radio frequency and op-

tical) as well as the need to manage blackout periods. NASA’s

DSN emerged as a viable option due to its proven heritage

with deep space communications. The key variables in the

trade space regarding radio frequency for data transmission

were gain, antenna diameter, data rates, mass, and heritage.

The mechanical design of the spacecraft involved researching

architectures from past missions. To minimize spacecraft cost

and fuel consumption, the SUNFLOWER Observatory space-

craft architecture emphasized mass and volume minimization

compared to previous heliophysics architectures. The core

spacecraft bus design variables were relative launch loads,

volume, mass, relative cost, and proven heritage.

The attitude determination system was compared across pre-

cision, power required, mass, cost, and proven heritage. The

attitude control system was compared across torque perfor-

mance, lifetime, power required, mass, cost, and proven

heritage. The power system was compared across power

generation efﬁciency, heat dissipation effectiveness, radiation

degradation, mass, cost, and proven heritage.

The thermal systems were designed with constraints and

requirements deﬁned by the other subsystems within the

SUNFLOWER vehicle. A trade space was analyzed for the

thermal control of the spacecraft for the extreme hot case

of 0.8 AU from the Sun and the extreme cold case of 1.1

AU from the Sun. The limited cooling capacity of passive

radiators indicated that actively cooled systems would be

necessary given the environment in which the SUNFLOWER

Observatory mission will operate. The spacecraft will use a

combination of passive radiators and pumped ﬂuid loops to

maintain the vehicle temperature. Pumped ﬂuid loops have

proven ﬂight heritage and have ﬂown on other, similar solar

missions in the past [28]. For speciﬁc subcomponents that

may require more granular thermal control, cryocoolers may

be implemented on an as-needed basis.

Finally, the electrical power system (EPS) is the interstitial

tissue that binds the other subsystems of the bus and allows

them to function. Using the power input available at the

limiting orbital distance of 1.1 AU and available mass and

volume constraints, the SUNFLOWER EPS team calculated

a necessary solar panel area of 25.4m2to support 8.5kW of

power consumption during propulsive maneuvers and 0.6kW

when collecting heliophysics data. A summary of the power

budget is shown in Table 3.

F. Risks

Sixteen risks were identiﬁed across four subteams: SE, HSM,

bus design, and CDA. Table 4 enumerates these sixteen risks

in the NASA standard risk matrix, as they were presented at

the Preliminary Design Review (PDR) [29].

During the PDR phase, between one and ﬁve concurrent mit-

igation strategies were identiﬁed for each risk, with special

attention given to the three highest risks presented at CoDR:

2.02 Payload Radiation Management, 3.01 Bus Radiation

Shielding, and 3.06 Mass Budget Closure. Risks 2.02 and

3.01 were both reduced through further analysis leading to

a more reﬁned understanding of radiation effects on the

payload (2.02) and a more radiation-robust design (3.01). An

improved understanding of the relationship between launch

vehicle payload mass and constellation deployment time re-

leased the dry mass constraint on the spacecraft bus, leading

to the retirement of risk 3.06 at PDR. Risk 1.01 was discarded

for being redundant with two others (2.02 and 3.01).

Risk statements are available in Appendix B, Table 16. The

mitigation strategies that led to risk score changes are enu-

merated in Appendix B, Table 17, where the bolded mitiga-

tion strategies represent the actions that were taken en route

to PDR to reduce risk scores.

IV. RES ULTS

This section presents design recommendations and analysis

results for the scientiﬁc instruments, constellation, and bus, as

well as a notional CONOPS and cost and schedule estimates.

Table 1.SUNFLOWER Bus Trade Space

Component Option 1 Option 2 Option 3

Attitude Determination Star Tracker Sun Sensor Gyroscopes

Attitude Control Reaction Wheels Thrusters Sun Radiation Pressure Control

Power System Identical Solar Arrays Primary + Secondary Arrays RTGs

Communication Band S-band X-band Ka-band

Structure Rack & Rail “Box” Unibody

Size Small/Microsat Medium Class Large Class

Thermal Control Cryocoolers Fluid Loops Radiators

Table 2.SUNFLOWER Bus Design Constraints

Characteristic Value

Mission Lifetime 11 years

Orbital Environment 0.8–1.1 AU

Communication Data Rate 500 MB/day min.

Onboard Data Storage 4 GB min.

Propulsion Power 7231 W

Instrument Power 21.5 W

Operational Temp. Constraint 300 K

S/C Pointing Accuracy 0.16°

S/C Pointing Knowledge 0.08°

Maximum Wet Mass 1000 kg

Table 3.SUNFLOWER Power Budget

Subsystem Thrust Mode (W) Sci. Ops. (W)

Payload 0 22

Propulsion 7231 0

Comms 328 328

Other Bus 194 194

Margin (10%) 775 55

Mode Total 8528 599

Table 4.NASA Risk Matrix at SUNFLOWER PDR

32.01 4.01 3.01

24.02 2.02

3.03

2.03

4.03

LIKELIHOOD

11.02

3.05 3.04

3.02

3.08

4.05

1.03

4.04

1 2 3 4 5

CONSEQUENCE

A. Recommended Architecture: Instrument Selection

The spacecraft payload bay will consist of three instruments

that were selected to fulﬁll the overall science objectives of

supporting global heliophysics research and facilitating more

accurate space weather predictions. The Science Traceability

Matrix (STM) in Appendix C offers a full mapping of sci-

entiﬁc goals to chosen instruments. Some key properties of

the preliminary instruments selected for SUNFLOWER—the

white light Compact Coronagraph (CCOR-1) [30], [31], the

Compact Doppler Magnetograph (CDM) [18], and the Sun

Watcher using Active Pixel System Detector and Image Pro-

cessing (SWAP) EUV imager [32], [33]—are shown in Table

5. Included in this table are each instrument’s inner ﬁeld of

view (FOV) and outer FOV, both measured in arcminutes (’),

as well as signal-to-noise ratio (SNR). In the case of SWAP,

the SNR ranges from 18.3 (while observing a quiet Sun)

to 183 (while observing an active region) [33]. The SNR

for CCOR-1 could not be determined from limited publicly

available information but is expected to meet mission needs.

Table 5.Final Instruments and Speciﬁcations

Instrument Mass Power FOV SNR

CCOR-1 21 kg 12.2 W 59’–299’ - - -

CDM 16 kg 6.7 W 0’–40’ 4000

SWAP 11 kg 2.6 W 0’–54’ 18.3–183

The CDM will collect full-surface magnetograms to observe

CMEs, sunspots, and coronal holes. SWAP and CCOR-

1 will provide overlapping solar disk images in different

wavelengths (visible and ultraviolet) that will track the cre-

ation and journey of CMEs, solar wind, and the impacts of

coronal holes on solar wind. Collectively, these payloads

will observe CMEs, sunspots, and coronal holes continuously

across the entire surface of the Sun, including the poles.

These observations, in turn, will fulﬁll the speciﬁc scientiﬁc

objectives of determining the Sun’s magnetic ﬂux boundaries

and determining the primary drivers of solar wind.

Stakeholder conversations indicated that miniaturizing the

instrument payloads would enable a lightweight spacecraft

design without compromise to mission objectives. The ﬁnal

payload selections emphasized low-mass conﬁgurations of

similar payload technologies. Potential risks due to low

TRLs for these selections were mitigated on an instrument-

speciﬁc basis. For example, the ﬁnal chosen (miniaturized)

magnetograph instrument is developed by the Southwest Re-

search Institute and is at TRL 6 at time of writing [18]. The

chosen white light coronagraph is one developed by the Naval

Research Laboratory and is currently expected to be used

on the NOAA GOES-U mission in 2024 [17]. The chosen

EUV imager already has ﬂight history on ESA’s technology

demonstration mission PROBA-2 [16].

Note that each of these instruments is remote sensing in

nature, not in situ. This decision is a result of stakeholder

conversations between the HSM team and heliophysics com-

munity that recommended prioritizing novel vantage points

of remote sensing as a means of propelling the heliophysics

ﬁeld forward. As such, the mission is designed to prioritize

continuous polar coverage via remote sensing. Moreover,

these same stakeholder conversations indicated that a more

direct zenith viewing angle, deﬁned as ζin Figure 5, of the

solar surface—especially the poles—would generate higher-

quality remote sensing data. Requirement 4.02, Solar Pole

Viewing Angle, was created to capture this need and passed

to the CDA team for constellation design consideration.

Figure 5.Deﬁnition of the zenith viewing angle ζ.

For each instrument, there will be two stages of calibra-

tion. Pre-mission calibration, which will occur on the

ground in controlled laboratory environments, will be the

ﬁrst stage. The other stage of calibration will be performed

semi-regularly during mission operations. These standard

calibrations are performed for astronomy data—bias, dark,

and ﬂat corrections. Bias corrections (included in dark

and ﬂat corrections) account for image sensor noise, dark

corrections account for thermal effects on the instrument, and

ﬂat corrections account for dust and vignetting. The focus

will also be calibrated semi-regularly for each instrument to

ensure that data remains high quality. Additionally, the CDM

will require calibration to remove any small magnetic bias

from the instrument [18]. All of these calibrations can be

performed on the instrument, as the requisite operations, such

as subtraction and division, are minimally intensive.

The limiting factor on cadence for the SUNFLOWER mission

is the amount of data to be downlinked each day. Due to

the large distance from Earth to many of the satellites, the

data downlink budget will be smaller than the data budget

for other heliophysics missions that are closer to the DSN.

Additionally, due to the conﬁguration of the satellites in the

constellation and the layout of the DSN on Earth, there will

be satellites below the ecliptic that will have less DSN contact

time and a correspondingly lower amount of data downlinked

per day. The maximum science data downlink per day in

the worst-case scenario is 435.5 MB per day in the case of

satellites below the ecliptic at a distance of 2 AU from Earth.

For satellites at and above the ecliptic, this value rises to be

871.0 MB per day.

Each instrument on the satellite produces data ﬁles that will

be in the Flexible Image Transport System (FITS) image

format. The ﬁle sizes per instrument were determined from

actual instrument data (like SWAP [34]), data from similar

instruments (like CCOR-1 [35]), or from image dimensions

(like CDM [18]). Additionally, on-board compression is

needed, as speciﬁed by NASA’s parameters on FITS com-

pression [36]. The total and compressed ﬁle sizes are pre-

sented in Table 6.

Table 6.Instrument Data File Sizes

Instrument File Size Compressed Size

CCOR-1 2 MB 0.50 MB

CDM 7 MB 1.75 MB

SWAP 2 MB 0.50 MB

Using these ﬁle sizes, the cadence for each instrument was

determined for satellites both above and below the ecliptic.

The optimal cadences for below the ecliptic can be found in

Table 7 and those at or above the ecliptic in Table 8.

Table 7.Optimal Instrument Cadences Below Ecliptic

Instrument Cadence MB/Day Images/Day

CCOR-1 15.0 min 48 96

CDM 7.0 min 360 205

SWAP 30.0 min 24 48

Table 8.Optimal Instrument Cadence At/Above Ecliptic

Instrument Cadence MB/Day Images/Day

CCOR-1 7.5 min 96 192

CDM 3.5 min 720 410

SWAP 15.0 min 48 96

After compression, the images will be sent to Earth, where

calibration, post-processing, and data analysis will occur.

As mentioned in Section I, instruments with low annual sen-

sor degradation rates were selected to meet mission lifetime

requirements. Speciﬁcally, the imaging sensors of CCOR-1

and SWAP showcase longevity in imaging sensor degrada-

tion, degrading at only 0.45% and 0.35% total pixel failure

rates every year, respectively [12], [37]. Figure 6 shows the

expected degradation over the expected initial lifetime of 11

years of these imagers, which lose less than 5% of all pixels

after one solar cycle [13]. For comparison, one of SDO’s

ultraviolet imaging sensors degraded over 90% at the 105 nm

wavelength after just a few years [12].

Figure 6.SUNFLOWER payload sensor degradation.

B. Recommended Architecture: Constellation Astrodynamics

A trade study for the constellation design was performed

which resulted in the constellation shown in Figure 7, a clas-

sic 3-plane Walker-Delta 54.7◦constellation [38]. In addition

to a 120◦difference in the longitude of the ascending nodes,

each plane employs a ±80◦true anomaly offset relative to

the rear and forward planes, respectively, to optimize zenith

viewing angles across the solar surface.

Figure 7.Orbits for a 3-plane Walker-Delta

constellation with 54.7° inclination.

Evenly distributed across the planes, the nine satellites

achieve constant 4π-sr coverage of the solar surface with a

maximum zenith angle of 61◦globally and just 49◦at the

solar poles, as shown in Figure 8. It is worth noting that this

ﬁgure does not take into account differential solar rotation.

The sinusoidal variation across longitude is due to the 7◦tilt

between the solar equator and the ecliptic, as depicted by

Figure 9. The remaining undulating pattern results from the

particular conﬁguration of the orbital planes.

Figure 8.At the solar poles, the maximum zenith

viewing angle ζfor the constellation is 49◦.

The choice of direct injection to high-inclination orbits for

improved observation of the solar poles necessitates unusu-

ally high ∆v. While electric propulsion requires approxi-

mately 60% more ∆vto reach the ﬁnal orbit than chemical

propulsion, the high speciﬁc impulse (Isp ) from the tech-

nology offers substantially higher ∆vper unit propellant

mass. As a result, the recommended propulsion technology

that emerged from the trade study was NASA’s Evolutionary

Xenon Thruster-Commercial (NEXT-C) gridded ion engine

[39]. Figure 10 shows that the proposed propulsion solution

requires signiﬁcantly less propellant mass to achieve the same

∆vcompared to other available technologies.

Figure 9.The constellation orbits are deﬁned in the

ecliptic plane, tilted 7◦relative to the solar equator.

NASA recommended using the NEXT-C ion engine in its

2013 Planetary Science Decadal Survey [40]. As such, the

engine’s novel capability is a key enabler of this constellation.

Figure 10.Propellant mass required to achieve a

desired ∆vfor varying engines sorted by Isp. Assumes an

individual spacecraft dry mass of 419.6 kg, including 15.4

kg of hydrazine reserved for momentum desaturation.

The launch vehicle selected will also provide about 9.2 km/s

of the approximately 38.2 km/s ∆vrequired per spacecraft to

reach the ﬁnal orbits, with the rest from the NEXT-C gridded

ion engine on each spacecraft. From a survey of the avail-

able launch vehicles, considering cost, risk, and capability,

the CDA team selected the Falcon Heavy combined with

a STAR 48BV upper stage to launch three SUNFLOWER

Observatory spacecraft towards each Walker-Delta orbit (data

provided by [41], [42]).

On this basis, a minimum of three launches—one per or-

bital plane—will be necessary. Although an option exists

to increase the number of launches to accommodate one

spacecraft per launch vehicle, thereby decreasing the wet

mass and overall time of ﬂight for each spacecraft, exercising

this option will substantially increase the cost.

Furthermore, the solution is relatively insensitive to variations

in the ﬁnal spacecraft dry mass. An increase in mass increases

time of ﬂight rather than making the trajectory infeasible. As

the bus design team settled on the ﬁnal design, the CDA team

reﬁned the ∆vand trajectory estimates at a spacecraft dry

mass of 419.6 kg, including 15.4 kg of hydrazine reserved

for momentum desaturation. With an analytical thrust rule

for inclination change, and optimized phasing maneuvers, the

resulting ∆vbudget is given in Table 9.

Table 9.∆vBudget for a Time of Flight of 3.2 Years

Maneuver ∆v(km/s)

Inclination Raising 24.4

Phasing 2.6

Station-Keeping 0.1

Margin 1.9

NEXT-C Total 29.0

Launch Vehicle V∞9.2

The total expected ∆vfrom the NEXT-C xenon engine for

the direct injection, including inclination and phasing, is 27.0

km/s. The heliocentric science operations only require about

0.1 km/s for station-keeping over the entire mission lifetime.

This leaves about 1.9 km/s of ∆vas margin for a total of 29.0

km/s of onboard ∆vfrom the NEXT-C ion engine.

C. Recommended Architecture: Spacecraft Bus

Table 10 captures the recommended bus design decisions.

Table 10.SUNFLOWER Bus Design Decisions

Design Space Decision

Attitude Determination Star Trackers + Sun Sensors

Attitude Control Reaction Wheels + Thrusters

Power System Identical Solar Arrays

Communication Ka-band + X-band Backup

Structure Rack & Rail

Size Small/Micro

Material Machined Aluminum

Based on these design choices, Table 11 offers a mass budget

broken down by subsystem. The resulting spacecraft has a

dry mass of 404.2 kg (or 419.6 kg including hydrazine) and a

wet mass of 849.6 kg.

A labeled depiction of the stowed vehicle architecture is

presented in Figure 11. Aluminum brackets connect separate

levels of the chassis, each holding different subsystems.

During launch the vehicle is secured via hold down and

release mechanisms (HDRMs) on the rear face of the vehicle

opposite the parabolic antenna. This approach results in a ﬁrst

fundamental frequency of 40 Hz. Additionally, under random

vibration loading as prescribed by the SpaceX Payload User

Guide [43], the maximum deﬂection in this conﬁguration is 2

mm, which can be mitigated by adding ribs and supports. The

preliminary dynamics analysis performed for PDR showed

consistency with these requirements.

Once the vehicle is deployed from the kickstage, its solar

panels extend to begin powering the vehicle. The deployed

conﬁguration is shown in Figure 12.

Table 11.Spacecraft Estimated Mass Budget

Subsystem Mass (kg)

Payload 48.0

Structure & Mechanisms 62.0

Thermal 26.0

Power 75.4

Communications 29.5

ADCS 30.7

Propulsion 102.1

Margin 30.5

Total Dry Mass 404.2

Xenon Propellant 430.0

Hydrazine Propellant 15.4

Total Wet Mass 849.6

Figure 11.The integrated SUNFLOWER vehicle in the

stowed position.

Figure 12.The deployed SUNFLOWER spacecraft. The

volume of the vehicle is 1 m x 1 m x 3 m stowed, with 24.5

m2of solar panel area when deployed.

Given that the vehicle is in a heliocentric orbit that is virtually

always in direct sunlight, cooling the vehicle is critical.

Analysis of the trade space led to a combination of pumped

ﬂuid loops and passive radiators based on heritage from

previous solar missions and performance [28]. This analysis

showed that the system needs 15–20 kg of mass allocated

for a pumped ﬂuid loop system transferring heat load to the

radiative surface. The goal is for the ﬂuid loops to remove all

excess heat load from the electronics and transfer it onto the

radiative surface. With an operational temperature constraint

of 300 K, 26.5 m2of radiator area will be necessary. The rear

faces of the solar panels will supply most of this radiator area

to limit the need for additional deployable radiators.

During the conceptual design phase, intersatellite links (ISLs)

were investigated as a potential method for circumventing

communication blackout periods and enabling near-real-time

data transmission back to Earth. Two options were considered

for ISLs: optical terminals and K-Band antennas. A free

space optical satellite communications link budget was used

to calculate the required transmission power for a 1.73-

AU optical link [44]. Although the data rates and power

requirements for the optical ISLs were appealing, the pointing

requirements were orders of magnitude more challenging to

accommodate than for the Ka-band link to DSN (5.7∗10−5

degrees budgeted pointing error compared to 0.16 degrees).

Likewise, the link and data budget analyses for the K-band

ISL revealed a data transmission rate of only 1302 bps

at a distance of 1.73 AU. Moreover, Ka-band frequencies

experience minimal interference from solar radiation [45],

and the recommended constellation architecture enables all

SUNFLOWER SC to have direct line of sight to Earth during

nominal science operations, reducing the overall value pro-

vided by ISLs. In light of these facts, at PDR, ISLs were

removed in favor of backup X-band antennas for contingency

communication with DSN in case of Ka-band or ADCS

anomaly, and risk 3.07 Solar Conjunction was retired.

D. Concept of Operations

Prior to launch, three SUNFLOWER Observatory spacecraft

will be integrated into each of three Falcon Heavy fairings,

as depicted in Figure 13 [43]. Operations begin with three

Falcon Heavy launches from Earth, each approximately 60

days apart. Any error in launch window will directly affect

the phase of the ﬁnal science orbit with respect to Earth,

but the launch can be delayed by up to three weeks before

signiﬁcant degradation (>1%) of the zenith viewing angle

risks compromising compliance to Requirement 4.02.

Figure 13.Three SUNFLOWER vehicles shown in a

Falcon Heavy fairing on its side.

Each launch will ﬁrst target a parking orbit of 160 km in

altitude before ﬁring the kickstage to escape Earth’s orbit. A

brief stay in the low parking orbit enables the rocket to take

advantage of the high velocity to increase the escape V∞.

After the upper stage completes its burn, each of the three

spacecraft on a given launch will deploy from the rideshare.

The ion propulsion system will then deliver each spacecraft

to its ﬁnal inclination and phase over a period of 3.2 years.

During this transfer, the SC will change thrust direction twice

per revolution to maximize the fuel efﬁciency of the inclina-

tion change maneuvers. Given the low-thrust maneuvering,

there is ample time to turn off the engine and power on the

instruments to collect data or calibrate instruments during the

transfer orbit without affecting time of ﬂight. This is a key

beneﬁt of the direct transfer, as the SC will always remain

between 0.8 AU and 1.1 AU (with phasing) from the Sun.

At the end of the transfer orbit, the spacecraft will power

down the ion engines and orient their science instruments

toward the Sun to begin science operations. They will deliver

continuous 4π-sr coverage of the Sun with nearly continuous

communication to Earth until decommissioning.

The decommissioning procedure will be highly dependent

on the context at the end of the eleven-year mission. If the

spacecraft are functional after the eleven years, extended op-

erational support can be requested by relevant stakeholders.

Should the operators choose to decommission, the ground

operations will command the spacecraft to disable battery

charging and transmission capabilities of the spacecraft. This

will ensure the spacecraft will not perform unplanned maneu-

vers or cause unnecessary radio interference as their batteries

drain one ﬁnal time. Further analysis regarding the stability

of the heliocentric orbits over a long time horizon after

decommissioning is recommended as future work.

E. Cost and Schedule

Figure 15 in Appendix D depicts a preliminary quarterly

schedule estimate for the SUNFLOWER program, beginning

in FY2025 and extending for roughly eighteen years through

planned decommissioning. A notional launch date of Dec.

31, 2031, in Q1FY2032 kicks off the operations phase of the

program, which continues until the notional decommission-

ing date of Dec. 31, 2042, at the end of Q1FY2043.

The NASA proposal for the FY2024 budget, at time of

writing, features a 6.7% reduced allocation for heliophysics

relative to the FY2023 budget [46]. Electing to begin the

SUNFLOWER program in FY2025 allows for the release of

the upcoming Decadal Survey for Solar and Space Physics

(Heliophysics) 2024–2033, which—it is hoped—will identify

an increased need for real-time space weather monitoring

and prediction. This would encourage NASA and the US

government to allocate additional funding to space programs

dedicated to heliophysics like the SUNFLOWER Observa-

tory, enabling the program to succeed.

The instrument development and delivery timeline leverages

the heritage of the SUNFLOWER instruments—which have

already launched [16], are manifested to launch in the near

future [17], or are at least space qualiﬁed and being prepared

for launch [18]—enabling an accelerated window of roughly

four years from a NASA Announcement of Opportunity to

instrument storage and preparation for delivery. Likewise,

the spacecraft and bus development timeline of four years

from kickoff through System Integration Review (SIR) is

derived from the corresponding timeline for the Ulysses

program [14]. The observatory integration phase following

SIR through liftoff, based on the corresponding timeline for

the Parker Solar Probe, is estimated to be roughly 1.5 years

[15]. Finally, the operations phase is baselined as eleven years

in accordance with the mission lifetime requirement.

Table 12 offers minimum, expected, and maximum cost

estimates, each with 20% margin, for the program over its

seven-year development timeline and eleven-year mission

lifetime. Each line item depicted in Table 12 represents

the summation of lower-level WBS line items omittedfor

brevity. One qualiﬁcation unit is included as part of design

and development costs, and the operations phase includes

both mission operation and data analysis costs.

In a departure from traditional program cost estimates, Table

12 allocates for the establishment of additional 34-m DSN

dishes at Goldstone, Madrid, and Canberra, as well as their

operation and maintenance during the SUNFLOWER mis-

sion. At time of writing, DSN supports about forty existing

deep space missions, with more than thirty additional projects

in development [47], [48]. With nine spacecraft in a con-

stellation demanding nearly continuous communication with

Earth, SUNFLOWER would place a signiﬁcant burden on the

limited number of 34-m dishes DSN is expected to have by

2031 [48]. Establishing additional stations at each DSN site

reduces this burden while simultaneously reducing recurring

engineering and mission operation costs for SUNFLOWER.

Table 12.Preliminary Cost Estimates

(in Thousands of FY2025 USD)

Min Exp Max

Design & Dev. 300,265 400,353 500,442

Build & Test 637,039 489,385 611,731

Launch & Kick 471,342 627,123 783,904

Operations 164,475 328,950 493,426

DSN Access 855,423 1,465,123 2,237,589

Margin (20%) 431,509 662,187 925,418

Total 2,589,053 3,973,121 5,552,509

The minimum, expected, and maximum cost estimates allo-

cate for the construction of three, ﬁve, and seven additional

34-m dishes at each site, respectively [23]. A cost compar-

ison shows that the establishment and maintenance of these

dedicated stations is comparable to—if not less expensive

than—the corresponding cost of nearly continuous access to

DSN for nine spacecraft over eleven years [49].

The total cost for the program is distributed by ﬁscal year

in Figure 14. Note that the annual costs never exceed the

FY2024 NASA Heliophysics budget request [46].

Figure 14.Total cost (in millions of FY2025 USD)

distributed by ﬁscal year.

The average lifecycle cost is estimated to be $4.0 billion in

FY2025 USD, which is substantial but ultimately consistent

with the number of spacecraft and launches and the scope

and lifetime of the program. Of this amount, roughly $2.1B

FY2025 USD captures non-recurring engineering—including

establishment of new DSN stations—while roughly $1.9B

FY2025 USD captures recurring engineering—including

DSN station operation and maintenance.

The average annual cost of SUNFLOWER during initial and

extended operations was compared to the same costs of the

GOES constellation [50], [51], the JPSS constellation [50],

[51], and the GPS constellation [52], [53]. The results

are given in Table 13. Overall, the average yearly cost of

SUNFLOWER—as low as $200 million in FY2025 USD for

mission extension—is comparable to that of NOAA Earth

observation and solar observation constellations like GOES

and JPSS.

The marginal value of SUNFLOWER forecasts for a “high

performance” event (like the 2003 Halloween storms) and a

“max performance” event (like the 1859 Carrington event) are

given in Table 14 [54]. This data is taken from a 2018 study

that examined the economic impacts of geomagnetic storms

in Europe, modeling current space-based forecasts against

improved forecasts beyond the Sun-Earth line [54]. Overall,

the improved forecasting offers marginal expected average

savings of $3.7B in global value yearly for a 2003-like solar

event and $32.5B yearly for a catastrophic Carrington-level

event [54]. Because this study assumed Europe to be the

primary target, these cost estimates would likely increase

for North America or Asia. Regardless, the exploration

of marginal value shows that SUNFLOWER would provide

value well beyond its costs in the event of intense solar

weather events, in addition to the substantial scientiﬁc value

provided by the constellation’s continuous observation of the

Sun’s entire surface.

V. CONCLUSION

A. Summary

The objective of the SUNFLOWER Observatory is to provide

continuous, comprehensive, long-term remote sensing across

the entirety of the solar surface. With the selected science

instruments, SUNFLOWER will provide heliophysicists with

unprecedented observational capability, enabling improved

understanding of the solar dynamo and more accurate space

weather predictions, among other beneﬁts.

The SUNFLOWER Observatory mission consists of nine

total spacecraft arranged in a three-plane heliocentric Walker-

Delta constellation. Each orbital plane contains three space-

craft in 1 AU circular orbits inclined at 54.7° relative to

the ecliptic plane. Critically, this constellation geometry

will allow for constant visual coverage of all solar latitudes

with a viewing angle of no more than 61° measured from

nadir. Spacecraft will be launched in groups of three onboard

Falcon Heavy rockets with STAR 48BV kickstages. Upon

separating from the kickstage, the spacecraft will use onboard

electric propulsion to phase apart and continually raise orbital

inclination until the target orbits are reached approximately

3.2 years after launch.

All nine spacecraft will be equipped with a full-disk Doppler

magnetograph, EUV imager, and white light coronograph, al-

lowing for observation of CMEs, sunspots, and coronal holes.

To satisfy the identiﬁed deployment and station-keeping re-

quirements, each spacecraft is equipped with a NEXT-C

gridded ion engine with 430 kg of onboard xenon propellant

as well as momentum-ofﬂoading thrusters with 15.4 kg of

hydrazine propellant. Further, each spacecraft will house

onboard star trackers, sun sensors, reaction wheels, twin solar

arrays with secondary batteries, a machined aluminum rack

and rail structure, and Ka-band nominal communication with

Earth via the Deep Space Network. Such a design results in

an estimated spacecraft dry mass of approximately 404.2 kg.

In the current schedule estimate, the mission’s instrument

and spacecraft bus development will start in FY2025 with

launches in FY 2032 and decommissioning in FY2043. SUN-

FLOWER will provide at least 7.8 years of continuous solar

observation, with the potential to extend the mission much

longer. With the present spacecraft design, this schedule

results in a total cost estimate of $4.0B in FY2025 dollars.

B. Future Work Beyond PDR

For future work, in the near term, the bus design team and

CDA architectures would be matured to higher ﬁdelity. For

the spacecraft bus, speciﬁc vendors for spacecraft compo-

nents and construction would be selected, and the thermal and

Table 13.Lifecycle Cost Comparison (in Millions of FY2025 USD)

GOES JPSS GPS SUNFLOWER

Lifecycle Cost 11,700 11,322 5,207 3,980

Average Cost per Year 334 808 158 200–362

Est. Value Added per Year 5,500 5,500 125,000 1,882–6,405

Table 14.Marginal Value Scenarios (in Millions of FY2025 USD)

High Performance Scenario Max Performance Scenario

Event Type 2003 Halloween Storm Carrington Event

Total Global Marginal Beneﬁt $37,635 $960,793

Likelihood/Timespan 20 yrs 150 yrs

Average Marginal Value over Timespan $1,882/yr $6,405/yr

structural analyses would be performed with industry-grade

software to increase conﬁdence in the proposed solution.

Likewise, the CDA team would work on nonlinear optimiza-

tion of the spacecraft trajectories to their ﬁnal science orbits

and a stability analysis of the various proposed heliocentric

orbits after decommissioning. Over a longer term, the project

would seek approval and funding through a NASA proposal.

APPENDICES

A. REQUIREMENTS MATRIX

Table 15 lists the set of requirements for the SUNFLOWER

program, including requirement identiﬁcation number, title,

description, rationale, category, and veriﬁcation and valida-

tion (V&V) method at PDR.

B. RISK MATR IX

Table 16 provides the title of each risk presented at PDR,

along with the corresponding risk realization in the form

of an if-then statement. Table 17 enumerates the potential

mitigation strategies for reducing the total risk score for

each risk through either likelihood or consequence reduction.

Mitigation strategies that were implemented during the pre-

liminary design phase are presented in bold.

C. SCIENCE TRACE ABI LIT Y MATR IX

The STM (Table 18) was developed to streamline the SUN-

FLOWER instrumentation suite. The STM also helps to

guide cadence determination and instrument suite calibra-

tion. The scientiﬁc questions are derived from current key

problems within both the space weather and heliophysics

spheres. Instrument requirements were derived from expert

assessments of relevant phenomena to ensure the collection of

the correct data [5] and from the performance speciﬁcations

of comparable instruments to ensure that the SUNFLOWER

suite would be competitive [55], [56]. Projected instrument

performance is derived from [30] for CCOR-1, from [18] for

CDM, and from [57] for SWAP.

D. GAN TT CHA RT

Figure 15 offers a quarterly Gantt chart of the SUNFLOWER

program. The schedule begins in FY2025 and is truncated

after FY2035 for easier viewing. Activities beyond the cutoff

are continued primary operations and decommissioning in the

73rd quarter, after the eleven-year primary mission.

ACKNOWLEDGMENTS

The SUNFLOWER Observatory team would like to thank Dr.

Nicola Fox and Dr. Jeffrey Hayes of NASA, for their review

of and feedback on the mission concept, and Dr. Rebecca

Thorndike-Breeze and Christine Casatelli, for sharing their

expertise in effective communication and presentation. The

Heliophysics and Solar Monitoring and Systems Engineering

teams would like to thank Dr. Sarah E. Gibson of the

National Center for Atmospheric Research for her wisdom

and suggestions with regards to instrument prioritization and

miniaturization, as well as the scientists and researchers at the

Center for Astrophysics |Harvard & Smithsonian for sharing

their wisdom in heliophysics and for providing the opportu-

nity for SUNFLOWER team members to acquire hands-on

experience with ground-based heliophysics monitoring tools.

Finally, the authors would like to thank the other student engi-

neers and scientists whose efforts matured the SUNFLOWER

mission over the course of several months, namely: Jared

Boyer, Yana Charoenboonvivat, Kevin Garcia, Angel Gomez,

Frank Gonzalez, Tahmid Jamal, Steven Liu, Helena McDon-

ald, John Pendergrast, Joseph Rowell, Nicholas Showalter,

Jackie Smith, Theodore St. Francis, Preston Tower, Christo-

pher Vargas, Sienna Williams, and Kathleen Xu.

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BIOGRAPHY[

Allan Shtofenmakher is the chief en-

gineer for the SUNFLOWER Observa-

tory program, as well as a PhD can-

didate at the Massachusetts Institute of

Technology (MIT). His research inter-

ests broadly reside at the intersection of

space situational awareness (SSA) and

control of multi-agent systems, with a

focus on tracking orbital debris using

in-space satellite sensors. He holds an

MS degree in Aeronautics and Astronautics from Stanford

University and a BS degree in Aerospace Engineering from

University of California, Irvine.

Daniel Gochenaur is a Constellation

Design and Astrodynamics engineer for

SUNFLOWER. He is a PhD student and

NDSEG fellow in MIT’s Department of

Aeronautics and Astronautics. Daniel

holds an MPhil in Supersonic Aero-

dynamics from the University of Cam-

bridge and a BS in Aeronautics and As-

tronautics from Purdue University. His

present research interests center on at-

mospheric entry and multidisciplinary design optimization.

Ben Waters is the systems engineer

on the SUNFLOWER Observatory mis-

sion primarily responsible for integra-

tion with the Heliophysics and Solar

Monitoring subteam. He is an SM can-

didate at MIT with interests in space sys-

tems focused on orbital debris remedia-

tion and risk management. He holds a

BS degree in Astronautical Engineering

from the U.S. Air Force Academy.

Robert Cato III is a systems engineer on

the SUNFLOWER Observatory mission

working closely with the spacecraft bus

design subteam. He graduated from MIT

in June 2023 with an SB in Aerospace

Engineering. He is a ﬂight systems en-

gineer at NASA JPL as of August 2023.

His interests include systems engineer-

ing, robotics, and autonomy.

Duncan Miller is the systems engineer

responsible for integration related to

the Technology Roadmapping subteam.

He brings 10 years of experience as a

naval aviator leading complex missions

involving remote sensing and reconnais-

sance. Duncan earned his BS in Physics

from the U.S. Naval Academy in 2012

and is pursuing his SM in Engineering

and Management from MIT.

Luke de Castro is a systems engineer

for the SUNFLOWER Observatory who

works closely with the Constellation De-

sign and Astrodynamics subteam. After

graduating with an SB from MIT in June

2023, he is pursing his SM in Aerospace

Engineering at MIT, studying optimal

trajectory design and control for agile

unmanned aerial vehicles.

Tai Zheng is a member of the Sys-

tems Engineering subteam for the SUN-

FLOWER Observatory mission. He

graduated from MIT in June 2023 with

an SB in Aerospace Engineering. His

career interests include space explo-

ration and systems engineering, and he

is a satellite systems engineer at Boeing

as of September 2023.

Alexander Koenig is the Constella-

tion Design and Astrodynamics lead for

SUNFLOWER. He is a graduate of MIT

(dual SB in Physics and Aerospace En-

gineering 2022, SM in Aeronautics and

Astronautics 2023) and presently works

as a Starlink GNC engineer at SpaceX

as of June 2023. Alongside constel-

lation design and operation, his pro-

fessional interests include observational

astrophysics, multi-messenger astronomy, and SSA.

Katelyn Sweeney is the engineer-

ing lead of the bus design subteam

for the SUNFLOWER Observatory pro-

gram. She specializes in dynamic and

thermal analyses of complex spacecraft

systems. She received her SB in Mechan-

ical Engineering from MIT in 2018 as

well as her MS in Engineering Sciences

and her MBA from Harvard University

in 2023.

Joel Jurado Diaz is the lead telecom-

munications engineer on the bus design

subteam for the SUNFLOWER Observa-

tory mission, as well as a visiting student

in the Department of Aeronautics and

Astronautics at MIT. He graduated from

Universitat Polit`

ecnica de Catalunya in

July 2023 with a BS in Telecommunica-

tions Engineering. His interests include

space communications, the architecture

of connectivity systems, and optical communications.

Alexis Lepe is the Heliophysics and

Solar Monitoring subteam lead. They

graduated with an SB in Aerospace

Engineering from MIT in June 2023.

Their interests include optical astron-

omy, spacecraft controls, and embedded

systems.

Claire McLellan-Cassivi is a member of

the Heliophysics and Solar Monitoring

subteam for the SUNFLOWER Observa-

tory mission. They are working on a dual

SB in Earth, Atmospheric, and Planetary

Sciences and Engineering, with plans

to graduate in February 2024. Their

interests include optical and x-ray as-

tronomy, communication systems, and

remote sensing systems.

Frederick Ajisafe is a member of the

SUNFLOWER Heliophysics and Solar

Monitoring subteam. He graduated

from MIT in June 2023 with an SB in

Aerospace Engineering, with plans to

complete an SM in Aeronautics and As-

tronautics. His interests include orbital

dynamics, data analysis, and applica-

tions of machine learning to systems en-

gineering and natural phenomena.

Akila Saravanan is a member of the

SUNFLOWER testbed subteam. She

graduated in June 2023 from MIT with

a dual SB in Aerospace Engineering and

Computer Science. She is continuing at

MIT in pursuit of an SM in Aeronau-

tics and Astronautics, working on opti-

mization problems to dynamically allo-

cate and schedule UAVs for search-and-

rescue operations. Her interests include

autonomy and perception.

Joana Nikolova is the primary ADCS

engineer on the bus design subteam for

the SUNFLOWER Observatory mission.

She graduated from MIT with an SB

in Aerospace Engineering in the spring

of 2023. She is currently working on

her SM in Aeronautics and Astronautics

at MIT, investigating machine learning

capabilities for scheduling ground based

observation of orbiting objects. Her

interests include control, dynamics, planning, and autonomy.

Clara Ziran Ma is a member of the

Technology Roadmapping subteam for

SUNFLOWER. She is an SM student in

the Technology and Policy Program and

Department of Aeronautics and Astro-

nautics at MIT. In her research, she fore-

casts the evolution of the launch industry

in the coming decades and uses atmo-

spheric chemistry simulations to model

rocket emissions. Clara has a BS in Sci-

ence, Technology, and International Affairs from Georgetown

University.

Leilani Trautman is a graduate teach-

ing assistant for the MIT space systems

engineering capstone course and serves

as an advisor to the SUNFLOWER team.

She is currently an MEng student in the

department of Electrical Engineering

and Computer Science at MIT, and her

research is focused on mission automa-

tion for the Mars Perseverance rover

through a collaboration with NASA JPL.

She earned her SB in Electrical Engineering and Computer

Science from MIT in 2021.

Nadia Khan is a graduate researcher in

the Engineering Systems Lab in MIT’s

Department of Aeronautics and Astro-

nautics, pursuing an SM in Technology

and Policy at MIT’s Institute for Data

Systems and Society. She served as a

graduate teaching assistant for the MIT

spacecraft systems engineering capstone

course. Nadia also led the heliophysics

case study for the NASA-MIT Advanced

Space Technology Roadmap Architecture (ASTRA) project.

Nadia is currently working as a business development trainee

in ESA’s Human and Robotic Exploration directorate.

Olivier de Weck is the Apollo Program

Professor of Astronautics at the Mas-

sachusetts Institute of Technology where

he is the director of the Engineering

Systems Laboratory. His research is

in systems engineering with a focus on

how complex technological systems are

designed and how they evolve over time.

He is a Fellow of INCOSE and a Fellow

of AIAA and serves as editor-in-chief of

the Journal of Spacecraft and Rockets.

Edward Crawley is the Ford Profes-

sor of Engineering at MIT. His research

has focused on architecture, design, and

decision support and optimization in

complex technical systems subject to

economic and stakeholder constraints.

From 2011 to 2016, he served as the

founding president of the Skolkovo Insti-

tute of Science and Technology, Moscow.

Prior to that, he served as the director

of the Bernard M. Gordon-MIT Engineering Leadership Pro-

gram. From 2003 to 2006, he served as the executive director

of the Cambridge-MIT Institute.

Table 15.Requirements Matrix

ID Requirement Title Requirement Description Rationale Category V&V

1.01 Spacecraft Launch Ve-

hicle Compatibility

Each spacecraft shall be physically compatible with the

SpaceX Falcon Heavy rocket and STAR 48BV motor

Outputs of launch vehicle trade study

analysis, offering optimal performance

weighed against cost and risk

Interface Inspection

1.02 Spacecraft Communi-

cation Network Com-

patibility

Each spacecraft shall be compatible with the NASA

Deep Space Network (DSN) for communication and

ranging

Pre-existing DSN infrastructure optimal

for deep space communication

Interface Inspection

1.03 Spacecraft Lifetime The suite of scientiﬁc instruments and spacecraft hard-

ware shall be designed for survival over 11 years in the

environment of the heliosphere

Average duration of a solar cycle is 11

years

Environment/

Reliability

Similarity/

Analysis

2.01 Scientiﬁc Instruments The suite of scientiﬁc instruments on each spacecraft

shall be capable of collecting information to support

global heliophysics research and provide more accurate

space weather forecasting

Flow-down from SUNFLOWER Obser-

vatory mission description

Functional Inspection

3.01 Spacecraft Attitude

Control Capability

Each spacecraft shall support a pointing performance

at least as accurate as 0.16° during communication,

science, and thrust operations

Required to support Ka-band commu-

nication with DSN at 2 AU, which is

limiting case for pointing performance

Performance Analysis

3.02 Spacecraft Attitude De-

termination Capability

Each spacecraft shall support a pointing knowledge at

least as accurate as 0.08° during communication, sci-

ence, and thrust operations

Required to support Ka-band commu-

nication with DSN at 2 AU, which is

limiting case for pointing performance

Performance Analysis

3.03 Spacecraft Data Com-

munication

Each spacecraft shall support transmission of at least

500 MB/day of state-of-health and science data to Earth

during nominal cruise operations

Flow-down from key scienceobjectives,

based on bus and science instrument

data generation analysis

Performance Analysis

3.04 Spacecraft Onboard

Data Storage

Each spacecraft shall be capable of storing at least 4 GB

of pertinent state-of-health and science data in the event

of a communication blackout

Flow-down from key stakeholders; en-

ables storage of 1 week of pertinent data

for transmission at a later date

Performance Analysis

4.01 Solar Surface Coverage The constellation of spacecraft shall have 4π-sr con-

tinuous observation of the Sun during nominal science

operations

Flow-down from SUNFLOWER Obser-

vatory mission description

Performance Analysis

4.02 Solar Pole Viewing An-

gle

The constellation of spacecraft shall have a zenith view-

ing angle ζof the Sun’s poles not to exceed 60° during

nominal science operations

Flow-down from key stakeholders Performance Analysis

4.03 ∆vfor Constellation

Placement

Each spacecraft shall be capable of providing at least

27.1 km/s of ∆vto support orbital inclination changing,

phasing, and station-keeping maneuvers after launch

vehicle separation

Corollary of Requirement 1.01, en-

abling greater ﬂexibility for meeting

overall mission objectives

Performance Analysis

Table 16.SUNFLOWER Risk Statements

Risk ID Risk Title Realization

1.01 Radiation Failure (REMOVED) IF the spacecraft modules are not sufﬁciently resistant to radiation failures,

THEN the modules could degrade in performance or fail completely

1.02 Fuel Requirements Prohibitive IF the amount of fuel required for each spacecraft is prohibitively high,

THEN the constellation architecture solution may not close

1.03 High Launch Vehicle Costs IF specialized launch vehicle (LV) performance or multiple launches are required,

THEN launch costs may become prohibitively expensive

2.01 Payload Thermal Management IF payload temperatures exceed maximum non-operational survival limits,

THEN payload functionality may degrade or cease

2.02 Payload Radiation Management

IF payload sensors, lenses, etc., are not properly protected for the heliosphere

radiation environment,

THEN instruments may degrade and data quality may decrease

2.03 Payload Pointing Degradation IF attitude control is degraded or lost due to hardware degradation or a C&DH error,

THEN payload observation capability may diminish or cease

3.01 Bus Radiation Shielding IF the SC does not have sufﬁcient radiation shielding or shielded components,

THEN data will be corrupted and components will fail or lose accuracy

3.02 Bus Thermal Management IF the SC has insufﬁcient thermal ofﬂoad or balancing,

THEN component functionality will break down due to unsafe thermal ranges

3.03 Bus Power Supply Failure IF the battery pack fails,

THEN the spacecraft may need to operate on photovoltaic power alone

3.04 Actuator Degradation/Failure

IF the actuators become incapacitated either through a procedural error or due to

long-term use,

THEN the spacecraft may suffer degraded attitude control capability

3.05 Link Budget Closure IF the spacecraft link budget can only support a limited data rate,

THEN the spacecraft may not be able to downlink all science data in near real time

3.06 Mass Budget Closure (REMOVED)

IF the spacecraft wet mass exceeds 666 kg,

THEN three spacecraft cannot ﬁt into a single LV, resulting in the need

for more launches and thus increased cost

3.07 Solar Conjunction (REMOVED) IF the spacecraft attempts to transmit data past the Sun to Earth,

THEN the data will lose coherency due to the noise produced by the Sun

3.08 Spacecraft Center of Mass and

Thrust Vector Alignment

IF the thrust vector is excessively misaligned with respect to the center of mass of

the spacecraft,

THEN there will not be enough hydrazine to desaturate the reaction wheels over the

mission lifetime

4.01 Missed Launch Window IF the SC or LV experiences delays that result in a missed launch window,

THEN the program will experience substantial schedule delays

4.02 Suboptimal Mission Design Tools

IF analysis and orbit optimization tools fail to offer full coverage of the orbit

tradespace or take too long to run,

THEN the constellation design may be suboptimal and could cause program delays

4.03 Excessive Time to Final Orbit IF the selected trajectory takes a long time to reach the ﬁnal desired science orbits,

THEN the mission will have a reduced timeline for generation of useful science data

4.04 Excessive Fuel Demands

IF the total mass and ∆vrequirements for the constellation exceed the capabilities

of current launchers,

THEN the mission concept could be declared infeasible due to inability to launch

under existing LV mass limits

4.05 Low TRL of NEXT-C Ion Thruster

IF the NEXT-C ion thruster is not available to achieve the required ∆vor high duty

cycle,

THEN the direct injection architecture may be infeasible

Table 17.Risk Mitigation Strategies

Risk ID Mitigation Strategies

1.01

1) Select radiation-hardened modules for use in spacecraft;

2) Build in additional aluminum shielding to protect from ionizing radiation;

3) Design architecture to be fully hardware redundant

1.02

1) Explore alternative propulsion options, such as electric;

2) Explore more efﬁcient trajectories and planetary ﬂybys;

3) Reduce redundancy within the spacecraft to save mass

1.03

1) Conduct thorough trade study on launch options to inform decision;

2) Consider leaner constellation architecture to minimize quantity of LVs needed;

3) Reduce SC mass to ﬁt multiple on one LV

2.01 1) Design thermal protection system with heritage techniques to meet payload temperature limits with margin;

2) Choose orbital distance from the Sun that reduces risk of exceeding non-operational temperatures

2.02 1) Radiation harden all relevant payload components;

2) Encase payload electronics in protective casing

2.03

1) Design the payload and science mission to require minimal maneuvering of each SC during science operations;

2) Design all attitude modes to allow scientiﬁc observation;

3) Consider phased array antenna to limit attitude adjustments needed for communications

3.01 1) Research and implement radiation hardening techniques and encoding requirements on non-robust components;

2) Include redundant components for critical modules

3.02

1) Research previous thermal balancing methods taken by similar solar observers and associated data;

2) Encode strict requirements on thermal management;

3) Add sufﬁcient heaters and radiators to improve thermal control

3.03 1) Perform a failure modes and effects analysis (FMEA) to determine what could cause battery failure;

2) Ensure there are at least two battery modules

3.04

1) Remove friction-based control mechanisms (e.g., reaction wheels);

2) Include sufﬁcient ∆vmargin for contingency attitude control maneuvers;

3) Include redundant attitude control systems in the design

3.05

1) Use larger, higher-gain antennas on spacecraft to increase data throughput;

2) Ensure sufﬁcient power is available for high-power RF output;

3) Add additional satellites to the constellation to decrease relay distance and increase data throughput;

4) Reduce amount of scientiﬁc data communicated back to Earth;

5) Conﬁrm there are no Ka-band conjunctions

3.06 1) Allocate less mass to scientiﬁc payload and scale the rest of the spacecraft accordingly

3.07 Early analysis showed this risk would not be realized either through ISLs or through a solar-noise-resilient antenna with

appropriate constellation architecture designed to enable line-of-sight communication directly to Earth

3.08 1) Store additional hydrazine;

2) Add a gimbal to the electric propulsion thruster to dynamically change the thrust vector

4.01 1) Build substantial margin into the schedule to reduce the likelihood of risk realization;

2) Design a constellation that allows for regular launch windows (e.g., no ﬂybys)

4.02

1) Build substantial margin into the schedule to reduce the likelihood of risk realization;

2) Maintain a rigorous schedule;

3) Rely on ﬁrst principles as a backup to novel optimization methods

4.03

1) Use gravity assists sparingly;

2) Investigate trades between trajectory time and propulsion technology selection;

3) Design a constellation that can generate useful science data on approach to ﬁnal trajectories/orbits

4.04

1) Maintain awareness of current LV capabilities when designing the constellation;

2) Investigate trades between SC mass and LV performance;

3) Explore possibility of launching smaller but more numerous spacecraft with distributed capabilities

4.05 1) Develop spare constellation architecture and trajectory with possibly worse performance but lower ∆vrequirements;

2) Track the NEXT-C engine development in current projects

Table 18.Science Traceability Matrix

Science

Goals

Science

Objectives

Scientiﬁc Meas. Requirements

Instr.

Instr. Performance

Physical

Parameters Observables Param. Goal Projection

What are

the ﬂux

boundaries

of the Sun’s

magnetic

sphere?

Examine coronal

holes to determine

structure and

location of ﬂux

boundaries

Magnetic

ﬂux around

coronal

holes, esp.

around poles

Magnetogram

at poles CDM AR

(arcsec) 1.0 1.2

Determine

occurrence of

active regions

Location and

number of

sunspots

Image of

photosphere SWAP

(arcsec) 1.00 3.17

What are

the primary

drivers of

solar wind?

max FOV

(SR) 3.0 2.5

Observe nature of

interaction

between CMEs

and coronal holes

in Sun’s magnetic

ﬁeld to determine

structure of

outgoing solar

wind

Magnetic

ﬂux and

structure of

corona

Coronagraphic

imaging CCOR-1

(arcsec) 6–30 33

FOV

(SR) 3.0–15.0 3.7–18.7

Continuous-

FOV EUV

imaging

SWAP

spectral

range

(nm)

17.1–19.3 16.6–19.5

LOS full-disk

photospheric

magnetograms

CDM AR

(arcsec) 7.0 1.2

Instr. = instrument, Param. = parameter, FOV = ﬁeld of view, LOS = line-of-sight, AR = angular resolution, SR = solar radii

Figure 15.Quarterly Gantt chart for the SUNFLOWER program.

Continuous In-Situ and Remote Sun Observation for Space Weather Monitoring and Mitigation of Infrastructure Threats Through an Optimized Heliocentric Satellite Constellation

Preprint

Nov 2024

Leonidas Askianakis

Although vital for life on Earth, solar activity poses questions and increasing threats to humanity due to the Sun's unknown dynamics, intensified by our dependence on terrestrial and space-based infrastructure. This situation is compounded by significant gaps in our understanding of space weather phenomena, the Sun's magnetic field, and the need for rapid responses to unpredicted solar events. To address these issues, an optimized heliocentric satellite constellation is proposed that leverages satellites in an Elliptical Walker Constellation. This system offers (among others) equally distributed arguments of periapsis separations and cross-coupled true anomalies with respect to the Sun-centric coordinate frame. In this paper it is also demonstrated that this strategic multi-spacecraft configuration makes it possible to distinguish spatial and temporal changes in solar wind phenomena, reconstruct, in 3D, Coronal Mass Ejections (CMEs), predict which space or ground-based infrastructure and when it will be affected by CMEs, maintain continuous coverage of the critical Sun-Earth line throughout the mission's duration, and protect future missions by providing simultaneously in-situ and remote measurements from small and cost-effective satellites.

Continuous in-situ and remote sun observation for space weather monitoring and mitigation of infrastructure threats through an optimized heliocentric satellite constellation

Article

Feb 2025

Leonidas Askianakis

A Review of the Extended EUV Corona Observed by the Sun Watcher with Active Pixels and Image Processing (SWAP) Instrument

Article

Full-text available

Oct 2022
SOL PHYS

The Sun Watcher with Active Pixels and Image Processing (SWAP) instrument onboard ESA’s PRoject for On Board Autonomy 2 (PROBA2) has provided the first uncompressed, high-cadence, continuous, large field-of-view observations of the extended extreme-ultraviolet (EUV) corona for over a complete solar cycle. It has helped shape our understanding of this previously understudied region, and pioneered research into the middle corona. In this article, we present a review of all publications that have utilized these observations to explore the extended EUV corona, highlighting the unique contributions made by SWAP. The review is broadly divided into three main sections of SWAP-based studies about: i) long-lived phenomena, such as streamers, pseudo-streamers, and coronal fans; ii) dynamic phenomena, such as eruptions, jets, EUV waves, and shocks; iii) coronal EUV emission generation. We also highlight SWAP’s imaging capabilities, techniques that have been applied to observations to enhance the off-limb observations and its legacy.

Economic Benefits of the Global Positioning System (GPS)

Technical Report

Full-text available

Jun 2019

This study quantifies the economic benefits of GPS for the United States' private sector to be

1.4 trillion (2017

) between 1984 and 2017. It also estimates that the economic losses from a disruption in GPS service would be $1 billion per day, and potentially more if the disruption were to occur during critical planting seasons for farmers. Economic sectors covered include precision agriculture, electricity, finance (high frequency trading), professional surveying, telematics, telecommunications (including 4G and 5G wireless), oil & gas, mining, maritime industries (e.g., recreational boating, port operations, commercial shipping, cruise lines), and location-based services (e.g., smartphone apps, mobile technologies).

Recent white-light coronagraphs at the Naval Research Laboratory

Article

Full-text available

Oct 2015

Russell A. Howard

A white-light externally occulted coronagraph is a telescope designed to view the very faint emission of the solar corona in the region close to the sun. It uses the Lyot principle to mitigate the effects of diffraction. The first such telescope flew on a spacecraft in 1971 and showed the dynamic nature of the corona. Since that mission, six other such coronagraphic telescopes have been flown, whose designs evolved to meet the requirements of the mission. This article describes the latest two coronagraphs and compares their capabilities and their designs.

On-Orbit Degradation of Solar Instruments

Article

Full-text available

Nov 2013

We present the lessons learned about the degradation observed in several space solar missions, based on contributions at the Workshop about On-Orbit Degradation of Solar and Space Weather Instruments that took place at the Solar Terrestrial Centre of Excellence (Royal Observatory of Belgium) in Brussels on 3 May 2012. The aim of this workshop was to open discussions related to the degradation observed in Sun-observing instruments exposed to the effects of the space environment. This article summarizes the various lessons learned and offers recommendations to reduce or correct expected degradation with the goal of increasing the useful lifespan of future and ongoing space missions.

The compact Doppler magnetograph (CDM) for solar polar missions and space weather research

Conference Paper

Aug 2022

Link Budget Analysis for Free-Space Optical Satellite Networks

Conference Paper

Jun 2022

NEXT-C Flight Ion System Status

Conference Paper

Aug 2020

Jack Fisher

The 7.4 kW NEXT-C (NASA’s Evolutionary Xenon Thruster – Commercial) gridded ion thruster system provides a combination of performance and spacecraft integration capabilities that make it uniquely suited for deep space robotic missions. With modifications, the NEXT-C system can meet some high total impulse defense and commercial missions as well. The purpose of the NEXT-C flight hardware development program, jointly funded by NASA and Aerojet Rocketdyne, was to establish a commercial supply of the thruster and power processing unit (PPU) for future NASA missions. The program has completed all development and flight production testing and delivered the first shipset of flight hardware to NASA GRC in early 2020 for use on the Applied Physics Laboratory’s DART (Double Asteroid Redirection Test) mission. The NEXT system was developed to a readiness approaching TRL 6 in the mid-2000s, followed by characterization and long duration testing at NASA Glenn Research Center. The original NEXT project effort culminated in a long duration life test of 50,000 hours on a NASA EM thruster using Aerojet Rocketdyne high fidelity optics. The thruster is throttleable across a thrust range of approximately 25 to 235mN. Thruster specific impulse ranges from 1400 to 4200sec, depending on the throttle condition. Each NEXT-C thruster is powered by a PPU with an input power of up to 7.4kW. The PPU converts spacecraft power, over an unregulated input voltage range from 80 to 160 volts, to the conditions required to operate the thruster, and also utilizes spacecraft 28 Volt power to operate the PPU’s communication and control circuitry. On the NEXT-C program, the component designs have matured to include design updates to increase capability, address issues identified during the development program, and incorporate lessons learned. Aerojet Rocketdyne has completed all program phases of the project, including full protoflight and integration testing of the Engineering Model hardware, Critical Design Review, ground test equipment validation, fabrication of flight hardware and protoflight level acceptance testing. Flight hardware testing included both component level testing of the flight PPU and thruster, culminating in an integrated system level hot fire test that demonstrated the flight controls and operating parameters planned to be utilized on the DART mission. The flight hardware testing was successfully completed to the DART mission requirements. This paper will present the flight system capabilities, latest test results, and flight hardware status for the NEXT-C ion engine system.

Quantifying the Economic Value of Space Weather Forecasting for Power Grids: An Exploratory Study

Article

Nov 2018
SPACE WEATHER

An accurate understanding of space weather socioeconomic impact is fundamental to the development of appropriate operational services, forecasting capabilities, and mitigation strategies. One way to approach this problem is by developing physics-based models and frameworks that can lead to a bottom-up estimate of risk and likely impact. Here we describe the development of a new framework to assess the economic impact of space weather on power distribution networks and the supply of electricity. In particular, we focus on the phenomenon of the geomagnetic substorm, which is relatively localized in time and space, and occurs multiple times with varying severity during a geomagnetic storm. The framework uses the AE index to characterize substorm severity, and the impact of the substorm is modulated by the resilience of the power grid and the nature of available forecast. Possible scenarios for substorm sequences during a 1-in-10-, a 1-in-30-, and a 1-in-100-year geomagnetic storm events are generated based on the 2003, 1989, and 1859 geomagnetic storms. Economic impact, including international spill over, can then be calculated using standard techniques, based on the duration and the geographical footprint of the power outage. Illustrative calculations are made for the European sector, for a variety of forecast and resilience scenarios. However, currently available data are highly regionally inhomogeneous, frustrating attempts to define an overall global economic impact at the present time.

U N I T E D Assessing the Economic & Social Benefits of NOAA Data

Article

Rodney Weiher

Satellite constellations

Article

Nov 1984
JBIS-J BRIT INTERPLA

J. G. Walker

The present investigation is concerned with the coverage of the earth's surface by multiple-satellite systems, taking into account results reported by Walker (1982). The considered studies have investigated systems with satellites following multiple circular orbits of equal period, providing continuous multiple (rather than only single) coverage of the entire surface of the earth. Attention is given to criteria regarding the selection of a satellite constellation, regular polyhedra and tesselations, practical satellite constellations related to delta patterns, sigma patterns, and omega patterns. Data concerning sigma patterns are presented in tables.