an engineering study - parker solar probeof a space mission, solar probe, to travel to the sun’s...

Solar Probe:An Engineering Study

Submitted to

NASA Goddard Space Flight CenterLiving With a Star Program Office

andNASA Headquarters

Sun-Earth Connection DivisionOffice of Space Science

November 12, 2002

Prepared by The Johns Hopkins University Applied Physics Laboratory,Laurel, Maryland, in partnership with the Jet Propulsion Laboratory

Work performed under Contract NAS5-01072

ii The Johns Hopkins University Applied Physics Laboratory

PREFACE AND ACKNOWLEDGMENTS

The NASA Solar Probe mission to the innerfrontier of the heliosphere is under study as apart of the Sun-Earth Connection theme withinthe Office of Space Science. This studywas conducted by the Johns Hopkins Univer-sity Applied Physics Laboratory (JHU/APL)

in partnership with the Jet Propulsion Labora-tory (JPL) under the sponsorship of theGoddard Space Flight Center Living With aStar Program Office. Participants and theirareas of contribution to this study arelisted below.

Participants* Areas of Contribution Roles in Solar Probe Study

Kenneth A. Potocki, Peter D. Bedini, NeilMurphy (JPL)

Program management Develop program schedule, costestimate

Ralph L. McNutt Jr., Barry J. Labonte,Bruce Tsurutani (JPL), Lawrence J.Zanetti

Study science Assure that mission conceptsatisfies science requirements

Douglas A. Eng, Douglas Lisman (JPL),Andrew W. Lewin, Laurence J. Frank

System engineering Develop Solar Probe concept

Yanping Guo, Robert W. Farquhar Mission design Define orbit, launch vehicle,propulsion requirements

Steven R. Vernon, Cliff E. Willey,Jennifer R. Tanzman, William E.Skullney

Mechanical engineering Design spacecraft structure;select materials

David W. Bennett, R. Kelly Frazer,Douglas S. Mehoke, FrankGiacobbe (SWALES)

Thermal engineering Design Heat Shield; selectmaterials

Carl S. Engelbrecht Propulsion engineering Design Propulsion System

Wade E. Radford, Per Uno Carlsson Power system engineering Design Power Subsystem

H. Brian Sequeira, Edward F. Prozeller,Robert S. Bokulic

Communications Develop Ka-band concept

Mark E. Pittelkau, Thomas E.Strikwerda, Steven J. Conard

Attitude Control Systemand Mission Design

Define attitude control,autonomous operation andpropulsion

Paul D. Schwartz, Martin E. Fraeman,Timothy S. Herder, Ark L. Lew, LloydA. Linstrom

Command and DataHandling

Design C&DH concept, flightsoftware

Paul H. Ostdiek, James E. Randolph(JPL)

Mission Engineering andTechnology

Develop technologies requiredto enable mission

Edward F. Prozeller, Kenneth P. Klaasen(JPL)

Instrumentaccommodation

Design Payload accommodationoptions

Raymond J. Harvey Mission operations Develop Integration andOperations

Pazhayannur K. Swaminathan Dust environment Develop concepts for spacecraftsurvival in dust environment

William N. Myers, Jia-Ning Huang,Cheryl L. Reed,

Fiscal and Administration Provide fiscal and administrativeleads

Mary Ann Haug Program administration Provide documentation andcommunications

Heara Lee, Stanley R. Purwin Jr. Quality assurance Provide quality assurance* Except as noted, participants are members of the JHU/APL staff.

Solar Probe: An Engineering Study iii

EXECUTIVE SUMMARY

Why is the Solar Probe study needed?

Humanity needs to understand how the con-nections between the Sun and Earth directlyaffect life and society. The first mission to theplanets, Mariner 2 in 1962, revealed a remark-able connection: the Earth is immersed in theoutermost material of the Sun, the solar wind.Since then, measurements of the propagationof mass and energy from the Sun to the Earthin the wind have indicated that the Sun is thedriving force affecting space weather. The dis-turbance of electrical power grids, communi-cation systems, aircraft and satellite systems byspace weather compels us to examine the ori-gins and sources of the solar wind in the mil-lion-degree solar corona with deeper physicalunderstanding.

The innermost heliosphere remains one of thelast unexplored regions of the Sun-Earth Con-nection (SEC). NASA recognizes the importanceof a space mission, Solar Probe, to travel to theSun’s corona in a voyage of exploration, dis-covery, and comprehension. The National Acad-emy of Sciences’ Space Studies Board 2002report, The Sun to the Earth—and Beyond, aDecadal Research Strategy in Solar and SpacePhysics, recommends that the Solar Probe beimplemented as soon as possible. The SECRoadmap 2002–2028 identifies Solar Probe asa mission to implement in the near term.

The Johns Hopkins University Applied PhysicsLaboratory (JHU/APL) was tasked to performa conceptual mission design study of the SolarProbe mission. This design study assesses thetechnical and programmatic feasibility of theselected mission concept to satisfy the SolarProbe science objectives. This report documentsthe engineering study. Cost methodology andestimates are provided to NASA in a separatedocument.

How was the Solar Probe study done?

The science investigation was derived from theNASA 1999 Science Definition Team Report,Solar Probe: First Mission to the Nearest Star.The fundamental science objectives of the So-lar Probe measurements are to understand theheating of the corona and the sources and accel-eration of the solar wind. Instrument accommo-dations were based on strawman instruments thatachieve the highest priority science objectives.

The Jet Propulsion Laboratory (JPL) has per-formed numerous Solar Probe mission designstudies over the last 20 years. To ensure thatthis study took full advantage of these past stud-ies, JHU/APL is teamed with JPL in this ef-fort. JPL has identified many options and tradespaces in the course of its studies. JHU/APLperformed a top-down look at the trade spaceissues, from documented science and solar en-vironmental requirements to engineering solu-tions, in order to propose a reliable concept tofit the mission.

What did the Solar Probe study find?

An efficient mission design and conservativeengineering spacecraft concept are proposed.The concept uses known materials, mature tech-nologies, and redundant design of critical sub-systems to enhance the probability of missionsuccess. The report describes a number of spe-cific design recommendations:

1. A mission trajectory has been designedwhich repeats every 13 months (baselinelaunch May 2010), has a 3.1-year cruise tothe first solar polar pass, and achieves a fi-nal 0.02- to 5-AU orbit about the Sun with aperiod of 4 years.

2. To measure solar wind variability during thesolar cycle, two solar passes form the

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baseline mission in 7.1 years; a third,optional pass can occur within 11.1 years.

3. A Jupiter Gravity Assist (JGA) trajectoryproduces a solar polar trajectory with 4-R

S

perihelion; an Atlas 551/Star 48B launchvehicle can provide the required liftcapability.

4. To improve instrument design and relaxspacecraft accommodation issues, the instru-ment power and mass resources are in-creased in comparison with the 1999 SolarProbe AO.

5. All instrument data will be acquired, re-corded, and transmitted continuously within0.5 AU (perihelion ± 10 days); additionalcruise-mode science data will be transmittedweekly throughout the mission during nor-mal housekeeping and navigation contacts.

6. A Ka-band downlink is used to minimizecoronal scintillation and meet data rate goalsnear perihelion; an X-band downlink sup-plies redundant backup capability. This datadownlink capability will augment the sciencereturn.

7. To protect the probe in the 3000-Sun solarintensity at perihelion, the thermal protec-tion system consists of a carbon–carbon coneprimary heat shield and a secondary heatshield with low-conductivity, low-massinsulation.

8. Three Multi-Mission Radioisotope ThermalGenerators (MMRTGs) will handle thepower and heating requirements reliablyduring the cold (Jupiter flyby) and hot

(Solar perihelion) phases of the mission andwill allow multiple flybys.

9. Redundant attitude control processors andsafehold control will assure that the space-craft bus stays within the umbra of thethermal protection system during the solarencounter.

Flying a mission to the Sun poses unique chal-lenges. The combined JHU/APL–JPL team hasinvestigated the technologies needed to conductthe proposed mission. A credible risk reductionplan, including fallback approaches for each ofthe key areas is described. This plan features aconcerted pre-Phase A and Phase A risk reduc-tion effort to mature key elements: the thermalprotection system material properties andmanufacturability; the characteristics of dustimpacts and accommodation in the spacecraftdesign; the efficiency of the Ka-band downlink;MMRTG development through the NASANuclear Initiative; and fail-safe attitude controlduring the solar flybys. Pursuit of this plan wouldallow overall programmatic and technicalrisk to be mitigated for mission requirementsdefinition.

How should the Solar Probe study beused?

Our society needs the scientific understandingthat Solar Probe will provide. Our scientific lead-ers advise that the science community is readyfor Solar Probe. This design study and its asso-ciated risk mitigation plan offer NASA a realis-tic path to implementing the Solar Probemission.

Mission SummaryLaunch opportunity every 13 months (Baseline launch May 2010)

Two solar passes (polar, 4RS) within 7.1 years; three within 11.1 yearsAtlas 551/Star 48B launch vehicle; 713 kg @ C3 = 128 km2/s2

JGA trajectory with post-perihelion �V for successive passes3.1-year cruise; 0.02 to 5 AU final orbit with period of 4 years

Measurement Strategy

• Characterize the solar wind within a high-speedstream

• Characterize the plasma in a closed coronalstructure and probe the subsonic solar wind

• Image the longitudinal structure of the white-lightcorona from the poles

• Produce high-resolution images in each availablewavelength

• Characterize plasma waves, turbulence, and/orshocks that cause coronal heating

• Determine the differences in solar wind character-istics during maximum and minimum solar activity

Flight System Concept• 15° half-angle conical carbon–carbon heat shield

• 3-axis stabilized with 0.2° pointing control and0.05° knowledge

• RTG power source (three Multi-Mission RTGssupply 330 W at beginning of life)

• Ka-band downlink, X-band uplink using 34-mDSMS dishes

• Data rate: up to 40 kbps real time at perihelion

• Data storage: redundant 128-Gbit stored data

• Instrument payload: 55 kg, 47 W

Technology Development

In situ instruments: Solar wind electrons and ion composition,magnetometer, energetic particle composition, plasma waves, and fast solar wind ion detector

Remote-sensing instruments: EUV imager, visible magnetograph–helioseismograph, and all-sky three-dimensional coronograph

Science Objectives

Category 1• Determine the acceleration processes and find the

source regions of the fast and slow solar wind atmaximum and minimum solar activity

• Locate the source and trace the flow of energy thatheats the corona

• Construct the three-dimensional coronal densityconfiguration from pole to pole and determine thesubsurface flow pattern, the structure of the polarmagnetic field, and their relationship with theoverlying corona

• Identify the acceleration mechanisms and locatethe source regions of energetic particles; deter-mine the role of plasma waves and turbulence inthe production of solar wind energetic particles

Category 2• Investigate dust rings and particulates in the near-

Sun environment• Determine the outflow of atoms from the Sun and

their relationship to the solar wind• Establish the relationship between remote

sensing, near-Earth observations at 1 AU, andplasma structures near the Sun

Category 3• Determine the role of x-ray microflares in the

dynamics of the corona• Probe nuclear processes near the solar surface

from measurements of the solar gamma rays andslow neutrons

Solar wind observations collected by the Ulyssesspacecraft during two separate polar orbits of the Sun,6 years apart, at nearly opposite times in the solar cycle.Near solar minimum (left) activity is focused at lowaltitudes, high-speed solar wind prevails, and magneticfields are dipolar. Near solar maximum (right), the solarwinds are slower and more chaotic, with fluctuatingmagnetic fields. (Courtesy of Southwet Research Instituteand the Ulysses/SWOOPS team)

NASA’s Sun-Earth Connection Theme seeks to under-stand our changing Sun and its effects on the solarsystem, life, and society.

The solar wind is the origin of magnetospheres and ofauroras on Earth, and Solar Probe will study the originof the solar wind.

The Solar Probe mission has been endorsed by theNational Academy of Sciences Decadal Survey andthe Sun-Earth Connection Theme Roadmap Team.

Solar Probe: An Engineering Study vii

TABLE OF CONTENTS

1. Mission History ...................................................................................................................... 1

2. Solar Probe Engineering Study Charter and Approach .......................................................... 2

3. Science Investigation .............................................................................................................. 33.1 Background .................................................................................................................... 33.2 Science Objectives ......................................................................................................... 43.3 Measurement Strategy .................................................................................................... 5

3.3.1 Requirements ...................................................................................................... 53.3.2 In Situ and Remote Sensing Instruments ............................................................ 63.3.3 Instrument Resource Allocations........................................................................ 6

3.4 Summary ........................................................................................................................ 7

4. Mission Implementation ......................................................................................................... 94.1 Mission Requirements .................................................................................................... 94.2 Mission Overview ........................................................................................................ 10

4.2.1 Mission Design Drivers .................................................................................... 104.2.2 Mission Design ................................................................................................. 104.2.3 Launch Vehicle Selection ................................................................................. 124.2.4 Mission Concept of Operations ........................................................................ 134.2.5 Mission Modes ................................................................................................. 144.2.6 Summary ........................................................................................................... 16

4.3 Spacecraft Overview .................................................................................................... 164.3.1 Spacecraft Concept Drivers .............................................................................. 164.3.2 System-Level Description ................................................................................ 19

4.4 Spacecraft Subsystem Descriptions ............................................................................. 234.4.1 Mechanical Description .................................................................................... 234.4.2 Thermal Design ................................................................................................ 274.4.3 Telecommunications Subsystem....................................................................... 344.4.4 Power ................................................................................................................ 384.4.5 Avionics Architecture ....................................................................................... 424.4.6 Data Handling Approach .................................................................................. 434.4.7 Guidance and Control ....................................................................................... 464.4.8 Propulsion ......................................................................................................... 504.4.9 Flight Software ................................................................................................. 51

4.5 Integration and Test (I&T) ........................................................................................... 524.5.1 Requirements .................................................................................................... 524.5.2 I&T Approach................................................................................................... 53

4.6 Ground and Data Systems Overview ........................................................................... 554.6.1 Requirements .................................................................................................... 554.6.2 Design Approach .............................................................................................. 554.6.3 Summary ........................................................................................................... 57

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4.7 Mission Operstions ....................................................................................................... 574.7.1 Requirements .................................................................................................... 574.7.2 Approach .......................................................................................................... 574.7.3 Summary ........................................................................................................... 59

5. Risk Mitigation ..................................................................................................................... 605.1 Thermal Protection System (TPS)................................................................................ 60

5.1.1 Primary Shield Assembly ................................................................................. 605.1.2 Secondary Shield Assembly ............................................................................. 635.1.3 Light Tubes ....................................................................................................... 64

5.2 Telecommuications....................................................................................................... 645.3 Dust Protection ............................................................................................................. 655.4 Attitude Control ........................................................................................................... 665.5 Radioisotope Power System ......................................................................................... 665.6 Risk Mitigation Activity Schedule ............................................................................... 67

Appendix A: Solar Probe Mass, Power, and Spacecraft Dimensions

Appendix B: References

Appendix C: Acronyms and Abbreviations

LIST OF FIGURES

Figure 1-1. The evolution of the Solar Probe spacecraft design through 1999. ........................ 1

Figure 3-1. Solar wind variation over the 11-year sunspot cycle. ............................................. 3

Figure 3-2. Structure in the solar wind. ..................................................................................... 4

Figure 3-3. Streamers are related to the slow solar wind. ......................................................... 5

Figure 4-1. Earth quadrature geometry for solar encounters. .................................................. 11

Figure 4-2. Solar Probe mission trajectory. ............................................................................. 11

Figure 4-3. Mission timeline showing optional third solar encounter andrelationships of the encounters to the solar cycle. ................................................ 12

Figure 4-4. Solar encounter trajectory details. ........................................................................ 12

Figure 4-5. DSMS coverage near perihelion. .......................................................................... 13

Figure 4-6. Solar Probe launch vehicle predicted lift capability. ............................................ 13

Figure 4-7. Operational concept summary. ............................................................................. 15

Figure 4-8. Solar Probe mission modes. ................................................................................. 15

Figure 4-9. Solar Probe external view deployed. .................................................................... 19

Figure 4-10. Physical block diagram of the Solar Probe concept. ............................................ 20

Solar Probe: An Engineering Study ix

Figure 4-11. Instrument mechanical accommodations. ............................................................. 24

Figure 4-12. Instrument fields of view. ..................................................................................... 24

Figure 4-13. Spacecraft subsystem mechanical accommodations. ........................................... 25

Figure 4-14. TPS mounted to spacecraft bus. ........................................................................... 28

Figure 4-15. Effect of cone half angle on primary heat shield temperature. ............................. 29

Figure 4-16. Effect of cone half angle on carbon mass loss rate. .............................................. 29

Figure 4-17. Thermal balance for the hot analysis case. ........................................................... 31

Figure 4-18. Hot case thermal analysis results. ......................................................................... 32

Figure 4-19. Cold case thermal analysis results. ....................................................................... 33

Figure 4-20. Comparison of Ka-band and X-band links during a solar encounter. .................. 35

Figure 4-21. Telecommunications architecture. ........................................................................ 36

Figure 4-22. Two-way, noncoherent navigation concept. ......................................................... 38

Figure 4-23. Power system block diagram showing one of the three MMRTGs. ..................... 40

Figure 4-24. Empirical RTG decay rate data. ............................................................................ 40

Figure 4-25. Solar encounter instrument stored data rates. ....................................................... 45

Figure 4-26. Nominal real-time science data collection profile. ............................................... 45

Figure 4-27. Solar encounter data collection and storage profile. ............................................. 46

Figure 4-28. Solar Probe coordinate frame. .............................................................................. 47

Figure 4-29. Solar Probe propulsion architecture. .................................................................... 51

Figure 4-30. Aerojet MR-111c 4.0-N rocket engine assembly, shown with axial nozzle. ........ 51

Figure 4-31. Spacecraft integration process. ............................................................................. 54

Figure 4-32. Solar Probe ground system architecture. .............................................................. 55

Figure 4-33. Timeline of a 6-month window around a solar encounter. ................................... 58

Figure 5-1. Risk mitigation schedule. ..................................................................................... 67

Figure A-1. Solar Probe Details. ............................................................................................ A-5

LIST OF TABLES

Table 3-1. Solar Probe sciece objectives. ................................................................................. 5

Table 3-2. The strawman science instruments used in planning instrument accommodation. 6

Table 3-3. Comparison of Solar Probe mission configurations. .............................................. 7

Table 3-4. Resource allocations to science instruments. ......................................................... 8

Table 4-1. Mission requirements derived from AO and SDT documenation. ......................... 9

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Table 4-2. Solar Probe key design drivers and resulting concept impacts. ............................ 17

Table 4-3. Solar Probe total radiation dose. ........................................................................... 18

Table 4-4. Hardware redundancy. .......................................................................................... 21

Table 4-5. Added functional redundancy in Solar Probe concept. ........................................ 22

Table 4-6. Solar Probe mass summary. .................................................................................. 22

Table 4-7. Power summary during solar encounter. .............................................................. 23

Table 4-8. Mechanical subsystem requirements summary. ................................................... 23

Table 4-9. Spacecraft temperature results of thermal analysis. ............................................. 33

Table 4-10. Telecommunication subsystem requirements summary. ...................................... 34

Table 4-11. Power subsystem requirements summary. ............................................................ 38

Table 4-12. Previous JHU/APL contributions to NEPA launch approval process. .................. 41

Table 4-13. Avionics subsystem architecture requirements summary. .................................... 42

Table 4-14. Mission Operations Team activities during mission phases. ................................ 58

Table 5-1. Subsystem components, heritage, and technology readiness levels (TRLs). ....... 61

Table A-1. Solar Probe Equipment List and Detailed Mass Breakdown.............................. A-1

Table A-2. Solar Probe Power Budget. ................................................................................. A-3

Solar Probe: An Engineering Study 1

1: Solar Probe Mission History

1. MISSION HISTORY

The idea of a Solar Probe mission dates to thefounding of NASA in 1958. The evolution ofthe mission concept can be thought of as a pro-gression of goals. The first goal, to reach solarorbit, led to the discovery of the solar wind byMariner II in 1962. The next goal, to sample thesolar system inside the orbit of Mercury, wasrealized in the Helios missions of the 1970s. Theability to launch spacecraft to Jupiter and usethe giant planet’s gravity field to assist in reach-ing other locations in the solar system, begin-ning with Pioneer 11’s flight to Saturn, openedopportunities for new types of solar missions.The goal of exploring high latitudes (Page 1975)was reached with Ulysses. Direct flight into thesolar corona, the source of the solar wind, be-came the Solar Probe concept.

The earliest studies of this concept were donein the mid-1970s, with the first major efforts

reported in the 1978 Jet Propulsion Laboratoryworkshops: A Close-up of the Sun (Neugebauerand Davies 1978).

The scale, science focus, and spacecraft configu-ration studied for Solar Probe have been revisedat intervals since that time. Initial suggestionsincluded dual spacecraft with different scienceinstruments. The mission was renamedStarProbe during the early 1980s, and becameFIRE as a joint United States–Russian missionin the mid-1990s. Figure 1-1 illustrates severalof the spacecraft configurations studied in de-tail and lists their basic elements. The two gen-eral concepts have either a conical primary heatshield with steerable high-gain antenna, or aparabolic heat shield/antenna combination. Nu-merous options have been considered for thespacecraft power source, bus configuration, in-strument suite, and mission plan.

Figure 1-1. The evolution of the Solar Probe spacecraft design through 1999.

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2: Solar Probe Charter and Guidelines

2. SOLAR PROBE ENGINEERING STUDY CHARTER AND APPROACH

2.1 Engineering Study Charter

The Johns Hopkins University/Applied PhysicsLaboratory (JHU/APL) was tasked to performa conceptual design study of the Solar Probemission. The study focuses on developing abaseline mission design concept for report to andconsideration by the Living With a Star Programat the NASA Goddard Space Flight Center(GSFC) and the Sun-Earth Connection Divisionat NASA Headquarters.

The study assesses the technical and program-matic feasibility of the selected mission conceptto satisfy the Solar Probe category 1 science ob-jectives. These science objectives were identi-fied in a 1999 report of the NASA ScienceDefinition Team, Solar Probe: First Mission tothe Nearest Star (Gloeckler et al. 1999). Costand schedule estimates for the mission conceptare developed and provided to NASA in a sepa-rate document.

2.2 Engineering Study Approach

The Jet Propulsion Laboratory (JPL) has per-formed numerous Solar Probe design studiesover the last two decades. To ensure that thisstudy took full advantage of these past studies,JHU/APL is teamed with JPL in this effort.

The study approach started with a JHU/APL top-down review of the science objectives and in-strument accommodation assumptions. Missionrequirements and a mission concept were de-veloped. The mission concept includes an or-bital trajectory to accomplish the sciencemeasurements, a candidate launch vehicle, anda concept of operations.

Using the mission concept and the space envi-ronments along the orbital trajectory, severalspacecraft concepts were developed. Subsystemtrade studies were conducted to evaluate everyaspect of these spacecraft concepts and to se-lect the best features for the proposed design.A series of reviews with scientists and engi-neers outside the study team were used to re-fine a reliable integrated spacecraft systemdesign.

Important considerations throughout the designare trade studies to characterize technical riskand establish a mitigation plan. The JHU/APLand JPL team has investigated the technologiesneeded to conduct the proposed mission. Beforeany technology was included in the baselinedesign, a credible risk mitigation plan was es-tablished, including fallback approaches for eachof the key technology areas.


3: Science Investigation

3. SCIENCE INVESTIGATION

3.1 Background

The primary science objective of Solar Probe is“to understand the processes that heat the solarcorona and produce the solar wind” (Solar ProbeScience Definition Team (SDT) final report,Solar Probe: First Mission to the Nearest Star,Gloekler et al. 1999). Since the discovery of thesolar wind by Mariner 2, NASA missions havetaken two complementary paths toward under-standing the origin of the solar wind. The firstpath measures the properties of the wind in situ.Sun-Earth Connection (SEC) missions on thispath are Voyager, which, at 86 astronomical units(AU) will soon reach the heliosphere’s outerboundary; Ulysses, which is studying the high-latitude heliosphere at 2 AU; and Wind and ACE,which are measuring the solar wind near theEarth at 1 AU. The second path uses remote sens-ing instruments to analyze the solar corona. SECmissions on this path are SOHO, with a diversecomplement of imaging and spectrographic in-struments; and TRACE, with very high spatialand temporal resolution imaging of corona struc-ture. These missions will be joined in 2004 byMESSENGER, which will measure the solarwind from Mercury orbit at 0.4 AU; and STE-REO, which will combine in situ measurementsand remote sensing from positions ahead of andbehind the Earth in its orbit.

Even so, the innermost heliosphere remains oneof the last unexplored regions of the solar sys-tem. The closest approach ever made to the Sun,0.31 AU (67 solar radii R

S), by the Helios space-

craft in the 1970s, is more than twice the outerlimit of any remote sensor. The smoothing ofsolar wind structures, caused by solar rotationand variations in propagation speeds, make de-tailed connections between the in situ and re-mote sensing measurements impossible. SolarProbe will make measurements in situ from thewind regime at 0.5 AU into the coronal volumesampled by remote sensing at 0.02 AU (4 R

S),

for the first time providing the physical connec-tions that will allow us to understand the solarprocesses that govern the solar corona and thesolar wind.

Past missions have revealed many aspects of so-lar wind acceleration (Figure 3-1), but compel-ling science questions remain unanswered.Coronal holes persist at the solar poles throughmuch of solar cycle. During sunspot minimum,the polar holes are present and the solar wind iswell organized, with the fast wind from the polarholes filling much of the heliosphere. The solar

Figure 3-1. Solar wind variation over the 11-yearsunspot cycle. Solar coronal images, from SOHO’sExtreme Ultraviolet Imaging Telescope (EIT), theMauna Loa coronameter, and SOHO’s Large Angleand Spectrometric Coronagraph (LASCO), show thecoronal structure at sunspot minimum (left) andmaximum (right). The speed of the solar wind, fromUlysses/SWOOPS, color-coded by the sign of themagnetic field, is plotted over latitude. At minimum,the solar magnetic field is dipolar, with large polarcoronal holes and streamers limited to the equator.The heliosphere is dominated by the fast solar windflowing from the polar holes. At maximum, the dipolarfield is gone, streamers are seen over a wide latituderange, and the wind contains a mixture of fast andslow wind intervals.



Figure 3-2. Structure in the solar wind. Structures observed out to 30 RS bySOHO/LASCO in polar solar corona holes have densities at least 4 timesthe background. No corresponding density structures are identified in thesolar wind measured by Ulysses/SWOOPS at 2 AU above the poles. SolarProbe will directly sample structure, from the wind into the corona, as it passesto a perihelion of 4 RS in its orbit (yellow line shows track of solar flyby).

magnetic fields restructure themselves during the11-year sunspot cycle. During sunspot maximum,the polar holes are absent, the solar wind is mixed,with fast and slow speed wind seen at all lati-tudes. As Figure 3-1 shows, the restructuring ofthe solar magnetic field, modifying the coronaand wind, gives important information about thegeneral sources of the fast and slow winds, butthe physical processes that accelerate the differ-ent wind speeds are not understood.

Observations from SOHO’s Ultraviolet Corona-graph Spectrometer (UVCS) imply that the windacceleration operates over the range of 2 to10 R

S in coronal holes. The wind flow appears

anticorrelated with structures in the coronalholes, as shown in Figure 3-2 (DeForest et al.2001). Low-density regions have higher tem-peratures, nonthermal, probably wave motions,and are possibly the source of the fast solar wind(Banjeree et al. 2000). How do the structures ofthe solar corona, so clearly correlated with theflow of the solar wind, evolve to the smootherwind observed at a distance?

The generation of the slow-speed wind withinthe streamers is a perplexing problem. SOHO/UVCS observations (Strachan et al. 2002) de-tected the wind outflow along the outer edgesand from the tops of streamers (Figure 3-3). Atthe same locations, discrete blobs of materialare seen moving outward in images from SOHO/LASCO images, but masses are too small tomatch the solar wind flux (Sheeley et al. 1997).Why does the wind from the streamers not ac-celerate to the faster speeds seen from coronalholes? Is the boundary between the closed andopen magnetic fields at the edges of streamersunstable to wave motions or discrete disconnec-tions that drive the blobs and remove energyfrom the wind flow?

3.2 Science Objectives

The specific science objectives that drove our en-gineering study were taken from the 1999 SolarProbe SDT final report (Gloekler 1999). No ef-fort was made to alter the objectives or prioritiesset by the SDT. The SDT prioritized the specific

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science objectives necessaryto achieve understanding ofthe processes that drive thesolar wind. The resultingprioritized list is given inTable 3-1.

We expect that a revisedSDT report must be pre-pared before any future So-lar Probe Announcement ofOpportunity (AO). Such areport would account foradvances since 1999 inscience understanding, in-strument technology, mis-sion resources, missionenvironment, and relatedspace missions. The SECRoadmap 2002 identifies anumber of missions that will



be contemporary with Solar Probe. In situ mea-surements of the inner heliosphere (0.3 to 1 AU)over all latitudes will come from the Solar–Terrestrial Probe (STP) Telemachus, the LivingWith a Star (LWS) Inner Heliospheric Sentinels,the European Space Agency (ESA) Solar Or-biter, and the joint ESA–Japan Bepi-Colombo.Remote sensing of structure in the solar atmo-sphere with order-of-magnitude improvementsin spatial resolution will come from Solar Or-biter and the STP Reconnection and Micro-scalemissions. Solar Probe science will be greatly

Table 3-1. Solar Probe science objectives.

Category 1 Objectives• Determine the acceleration processes and find the source regions of the fast and slow solar wind at

maximum and minimum solar activity.• Locate the source and trace the flow of energy that heats the corona.• Construct the three-dimensional coronal density configuration from pole to pole and determine the sub-

surface flow pattern, the structure of the polar magnetic field, and their relationship with the overlyingcorona.

• Identify the acceleration mechanisms and locate the source regions of energetic particles, and deter-mine the role of plasma waves and turbulence in the production of solar wind and energetic particles.

Category 2 Objectives• Investigate dust rings and particulates in the near-Sun environment.• Determine the outflow of atoms from the Sun and their relationship to the solar wind.• Establish the relationship between remote sensing, near-Earth observations at 1 AU, and plasma

structures near the Sun.Category 3 Objectives• Determine the role of x-ray microflares in the dynamics of the corona.• Probe nuclear processes near the solar surface from measurements of solar gamma rays and slow

neutrons.

enhanced by complemen-tary participation in thisensemble.

3.3 MeasurementStrategy

3.3.1 Requirements

The science objectives de-fine the basic mission de-sign. The orbital inclinationmust be 90° for passagethrough the polar coronalholes. The perihelion, overthe solar equator, must be4 R

S for passage over the

poles at 8 RS, where the fast

Figure 3-3. Streamers are related to the slow solar wind. Images forSOHO/UVCS in O5+ (red), neutral hydrogen (blue), and SOHO/LASCOelectron scattering (green) of a streamer. Contours of solar wind outflowspeeds are plotted at 0, 50, and 100 km/s, increasing outward. Blobs ofout-flowing material seen in LASCO movies accelerate along the sides ofstreamers in the locations show with arrows.

02-0817R-44

wind is still showing acceleration. Two perihe-lion passes are required at different sunspot cyclephases to adequately sample both fast and slowspeed wind regimes. Complementary remotesensing instruments are required to provide con-temporary context for the in situ measurementsthrough the unique range of solar latitudes anddistances traversed by the probe.

The SDT established a measurement strategyto accomplish the science objectives. The de-termination of the source regions and accelera-



tion processes of the solar wind depend on: (1)full knowledge of the plasma state, includingmagnetic field, composition, distribution func-tions, and wave spectra; and (2) knowledge ofthe context of the measurements from remotesensing of the local and global environment.Because the acceleration of the wind is likelyto be intimately connected with the heating ofthe corona, the same measurements are neededto understand the heating processes. Under-standing the role of wave–particle interactionsin heating and bulk acceleration processes re-quires rapid measurements of the dominant ionssimultaneous with plasma waves. Understand-ing the acceleration of energetic particles re-quires simultaneous measurement of thoseparticles (both ions and electrons) and of theplasma waves and properties of the dominantions at high cadence.

3.3.2 In situ and Remote SensingInstruments

Ideally, the instruments that make the requiredmeasurements would be well described andintegrated into a spacecraft design. However,to maintain fair competition, the proposalssubmitted in response to 1999 Solar Probe An-nouncement of Opportunity were embargoed.

For this engineering study, therefore, we basedinstrument accommodation on a set of strawmaninstruments that achieve the 1999 AO category 1Science Objectives (Table 3-2). For efficienciesin operation, we assumed that the instrumentswould be grouped into two suites. The in situinstrument suite comprises five instruments that,taken together, provide all of the in situ mea-surements needed to undertake the science ob-jectives. The remote sensing suite comprisesthree instruments that provide the context nec-essary to understand the in situ measurements.

Instruments designed to achieve Category 2 and3 science objectives may be proposed under afuture AO, pending resources and/or revisedpriorities. These could include a dust monitor, ahard x-ray and gamma ray spectrometer, and aneutron spectrometer.

3.3.3 Instrument Resource Allocations

The accommodation of the instruments with re-spect to fields of view and interference is de-scribed in Section 4. Several of the instrumentsrequire special accommodation. Because theVMH and EUVI image the solar disk, their viewmust be through light tubes that penetrate theprimary heat shield. The SWICES detects thebulk plasma. Because bulk flow is masked by

Table 3-2. The strawman science instruments used in planning instrument accommodation.

Suite Instrument Purpose1. Solar wind ion composition and elec-

tron spectrometer (SWICES)Measure distribution functions of the dominant chargestates of abundant ions and electrons.

2. Magnetometer (MAG) Measure vector DC magnetic field.3. Plasma wave sensor (PWS) Measure AC electric and magnetic fields.4. Fast solar wind ion detector (FSWID) Measure distribution functions of the most abundant

ions at high cadence.

In situ

5. Energetic particle composition spec-trometer (EPCS)

Measure fluxes of nonthermal electrons and abundantions.

1. All sky 3-D coronagraph imager (ASCI) Image the sunlight scattered from dust and electrons inthe volume around the spacecraft orbit to provide localcontext.

2. Visible magnetograph–helioseismo-graph (VMH)

Image the magnetic and velocity fields on the solarsurface to provide global context on the magnetic con-figuration through which the probe will pass.

Remotesensing

3. Extreme ultraviolet imager (EUVI) Image the thermal emission from the lower corona toprovide global context on the coronal topology throughwhich the probe will pass.



the heat shield during some parts of the probe’sorbit, a scoop is required to sample the plasmaduring those times.

The resources available to the science instru-ments are described in detail Section 4. Thesignificant differences between our study and themission described in the 1999 Solar Probe AOare summarized in Table 3-3. Specific differencesin resource allocations for the science instrumentsare shown in Table 3-4. As Table 3-4 shows, ourstudy resulted in more resources being madeavailable to the instruments. The additional re-

sources will help to mitigate instrumentdevelopment risk and will increase the scienceproductivity of the Solar Probe mission.

3.4 Summary

The scientific rationale for the Solar Probe mis-sion is well defined. The critical question iswhether the mission can actually be imple-mented as defined. Section 4 examines the vari-ous elements of the Solar Probe mission andspacecraft and defines a feasible and realisticconcept for the mission.

Table 3-3. Comparison of Solar Probe mission configurations. Major differences are highlighted in bold.

1999 AO Mission 2002 JHU/APL MissionMission Design

Launch vehicle EELV with Star-48V third stage EELV with Star-48B third stage(Atlas 551 used as baseline; compatible w/ Delta IV)

Launch date February, 2007 May 2010Pass 1 October, 2010 August 2013Gravity assists Jupiter — June 2008 (10.5 RJ) Jupiter — August 2011 (7.1 RJ)Pass 2 January, 2015 (Earth not in quadrature) August 2017 (Earth in quadrature)Radiation TID 88 krad (with RDM = 2) 90 krad (with RDM = 2)

Continuous for P – 10 through P +10 days (allpasses)

Data collection

Unavailable from P – 10 to P – 6 days,P + 3 to P + 10 days

Cruise mode science capabilitySpacecraft Design

3 MMRTGs (24 GPHS modules total; 330 W BOL)Power

1 RTG (18 GPHS modules; 300 W BOL)4.5 A-h lithium ion battery

15° conical carbon–carbon heat shieldParabolic carbon–carbon heat shieldAerogel (or carbon mat) secondary shield

Thermal Bus heating with RTG Bus heating with MMRTGsC&DH 6-Gbit data storage capability (redundant) 128-Gbit data storage capability (redundant)

3-axis stabilized 3-axis stabilizedIMU IMUStar trackers Star trackersSun sensors Sun sensors

Sixteen 4-N thrustersEight 0.9-N thrusters Reaction wheelsPointing accuracy (3σ) = 0.286° Pointing accuracy (3�) = 0.2ºPointing knowledge (3σ) = 0.057° (inertialhold)

Pointing knowledge (3�) = 0.057º

ACS

Pointing stability (3σ) = .0057º in 1 s Pointing stability (3�) = 0.0057ºSteerable 0.8-m HGA (dish on ~30° gimbal)Parabolic HGA

8-W Ka-band SSPA (downlink)8-W X-band SSPA (downlink) 8-W X-band SSPA (downlink backup)

Telecom Deep-space transponders X-band uplink, noncoherent navigation

Monopropellant hydrazine Monopropellant hydrazinePropulsion ∆V = 90 m/s ∆V = 150 m/s navigation, 50 m/s targeting burn



Table 3-4. Resource allocations to science instruments. Instrument re-sources and science return have increased in comparison with the 1999AO mission description.

Instrument Resources 1999 AOMission

CurrentStudy

Power (not to exceed values)Instruments 15 W 23 WData processing units (DPUs) andlow-voltage power converters (LVPCs) 24 W

Mass (not to exceed values)Instruments 21 kg 34 kgDPUs and LVPCsBoom and light tubes

14 kg 7 kg

Science return (data downlink capability)Real-time data rate (at perihelion) >5 kbps >25 kbpsOnboard memory 6 Gbit (�2) 128 Gbit (�2)Data return per perihelion pass 43 Gbits 128 to153 Gbits

Cruise-mode telemetry available Incidental benefit of noncoherent tracking for navigation One 8-hour contact per week provides 20 Mbits/week


4: Mission Implementation

4. MISSION IMPLEMENTATION

This section describes the general approach tothe implementation of the Solar Probe mission.The mission requirements are summarized, andoverviews of the mission design and spacecraftdesign are presented. Finally, the detailed descrip-tions of the spacecraft subsystems are presented.

4.1 Mission Requirements

The category 1 science objectives identified in1999 by the Solar Probe Science DefinitionTeam (SDT) (Gloekler et al. 1999) are the basisfor the current engineering study, and the mis-sion requirements derive directly from these ob-jectives. The instrument payload accommodatedby the mission and spacecraft designs is that

presented in the SDT report and described fur-ther in the Mission and Project Description(MPD) document (NASA 1999b) and the SolarProbe Announcement of Opportunity (AO)OSS-99-02 (NASA 1999a). Table 4-1 is a list-ing of the basic mission requirements.

The overarching objective of this study was todevelop a viable, affordable, detailed missionconcept that maximizes the science return. Ac-cordingly, in addition to the above mission re-quirements, the study goals included minimizingthe cost (without accepting undue risk), maxi-mizing the resources available for the instrumentpayload, and assuring maximum data downlinkthroughout the solar passes.

Table 4-1. Mission requirements derived from AO and SDT documentation.

Category No. Requirement Source1 Perihelion of 4 ± 0.1 solar radii (RS) MPD, p. 212 20-day solar encounter about perihelion MPD, p. 23 Provide instrument checkout and calibration soon after launch near 1

AUMPD, p. 2

4 Accurate post-mission trajectory reconstruction to a few kilometers MPD, p. 215 Two solar encounters MPD, p. 2

Mission

6 Orbit inclined 90° to the ecliptic MPD, p. 197 Accommodate all instruments described in AO MPD, p. 118 Provide light tubes and filter to limit flux to nadir-pointing imagers MPD, p. 259 Provide retractable boom to minimize interference with several of the

in situ instrumentsMPD, pp. 25,30, 49

10 Provide preflight purge to instruments MPD, p. 4911 Solar wind ion & electron spectrometer field of view (FOV) of

±20° boresighted to solar nadirMPD, p. 30

12 85° half angle FOV for spacecraft-mounted instruments MPD, p. 30

Mechanical

13 Accommodate at least 20.2 kg for instrument mass MPD, p. 1214 Provide regulated DC power to instruments between 22 and 36 V MPD, p. 2515 Accommodate at least 13.9 W average power for instrument power MPD, p 12

Power

16 Accommodate 100 W peak power; durations less than 50 ms MPD, p. 2517 Spacecraft temperature range of –20° to 50°C MPD, p. 28Thermal18 Spacecraft out-gassing ≤2.5 mg/s during solar encounters MPD, p. 4119 Provide 3-axis stabilized pointing during solar encounter and

instrument calibrationsMPD, p. 33

20 0.3° pointing accuracy to solar nadir boresight during the solarencounter

MPD, p. 34

21 0.05° inertial attitude knowledge during solar encounter MPD, p. 34

Attitude controlsystem (ACS)

22 0.005° in 1-s jitter relative to boresight during the solar encounter AO change listCommand & datahandling (C&DH)

23 Capability to handle peak data rates of 112.4 kbps from instrumentsuites during solar encounter

MPD, p. 12

Communications 24 Maximize real-time telemetry throughout the solar encounter MPD, p. 17



4.2 Mission Overview

This overview identifies the key drivers of themission concept and describes how the missiondesign accommodates them. The selection oflaunch vehicle and concept of operations is thendiscussed, along with a summary of the missionmodes.

4.2.1 Mission Design Drivers

Based on the above mission requirements andthe stated study goals, three fundamental fea-tures of the Solar Probe mission emerge as thekey mission drivers:

• Polar orbit with perihelion 4 RS from the cen-

ter of the Sun• Multiple passes of the Sun within one solar

cycle• Real-time data downlink during the passes

4-RS Polar Orbit. The science objectives require

that measurements be made at high solar lati-tudes, over the poles, and in the corona at a dis-tance of 4 R

S from the Sun’s center. Using

established methods of propulsion, the only wayto meet these requirements is to use a Jupitergravity assist trajectory. Supplying sufficientpower to the spacecraft at Jupiter (at 5 AU fromthe Sun) and within the orbit of Mercury (0.3AU) has not been accomplished with solar pan-els, but can easily be achieved by incorporatingtraditional nuclear power sources.

Associated with the 4-RS flyby is the need to har-

bor the instruments in the harsh solar environmentof the Sun’s corona. A thermal protection systemmust be capable of withstanding the extreme tem-peratures while rejecting enough heat to maintainthe spacecraft at normal operating temperatures.Precautions must be taken to account for the un-certain dust environment and to keep the space-craft shielded from the Sun as the spacecraft passesthrough perihelion at over 300 km/s.

Multiple Solar Passes. The science objectivesof Solar Probe as defined by the SDT require

sampling the solar wind in the corona at timesof high and relatively low solar activity, prefer-ably near solar maximum and minimum. Thisrequirement, coupled with the fact that the space-craft will spend a relatively short time in thecorona due to its great speed, led to the conclu-sion that at least two passes of the Sun shouldbe made. With the current mission design, So-lar Probe enters a 0.02 � 5 AU heliocentric or-bit that provides a first pass in 3.1 years, allowstwo passes in 7.1 years, and a third pass within11.1 years.

Real-time Downlink. Although the system de-sign is robust, and minimizing risk was a keyfactor in the various subsystem trade studiesperformed, the environment in the corona isharsh and uncertain, and there is some chancethat the spacecraft may encounter a catastrophicevent on any given pass of the Sun. As a result,another important driver is the capability to pro-vide real-time return of the science data. De-signing the mission for Earth to be in quadraturewhen the spacecraft is at perihelion creates theoptimum geometry to downlink the data (Fig-ure 4-1). Ka-band is used because coronal scin-tillations make such transmission difficult withX-band telecommunications.

4.2.2 Mission Design

The Solar Probe mission design uses a directJupiter gravity assist (JGA) to achieve a polarheliocentric orbit with the desired perihelion.The first solar pass occurs 3.1 years after launch,and a modest ∆V maneuver after this pass re-targets the spacecraft into a final orbit with so-lar passes possible every 4 years. The 4-yearperiod produces the same optimum communi-cations geometry for all perihelion passes. Threepasses are possible within an 11-year solar cycle.The mission profile used for this study assumesa launch in May 2010, but is available every 13months; Figure 4-2 shows the mission trajec-tory for the 2010 launch design; Figure 4-3shows a mission timeline.



View from Solar North Pole

(Not to Scale)

LOS

Approach Trajectory

Orbit Direction of Solar Dust

Sun

Earth

Solar Probe

at Perihelion

(4 RS )

~90°Solar Probe

at Beginning & End

of Encounter (0.5 AU)LOS

View from Solar North Pole

(Not to Scale)

LOS

Approach Trajectory

Orbit Direction of Solar Dust

Sun

Earth

Solar Probe

at Perihelion

)

~90°Solar Probe

at Beginning & End

of Encounter (0.5 AU)LOS

02-0817-4

Figure 4-1. Earth quadrature geometry for solar encounters.

from Cape Canaveral Air ForceStation (CCAFS), Florida.

Jupiter Flyby. The mission de-sign utilizes the gravity of Jupi-ter to lose enough angularmomentum to achieve a 4-R

S

perihelion. At the same time, thespacecraft makes a 90° planechange to enter into a polar he-liocentric orbit. Depending onthe launch date, the Jupiter flybyoccurs August 23 to 28, 2011.The closest approach to Jupiteris at 7 Jupiter radii (R

J) with a

flyby speed of 26 km/s. Asouthward Jupiter flyby ap-proach has been chosen ratherthan a northward approach be-cause the required launch energyis lower.

Sun

Jupiter

Earth

LaunchMay 26, 2010C3: 127 km2/s2

Solar Orbit 4 RS × 5.0 AU

Earth at Perihelion(at quadrature)

AphelionJuly 19, 2015

JGA FlybyAug 23, 2011C/A range: 7 RJ

1st Sun Encounter, Jul 18, 20132nd Sun Encounter, Jul 18, 2017

02-0817R-5

Figure 4-2. Solar Probe mission trajectory.

Launch. The 20-day launch window opens May26, 2010. The maximum required launch energyC

3 is 128 km2/s2, and the declination of launch

asymptote (DLA) ranges from –11.2° to –14.0°.The probe will be launched during the daytime

Solar Encounter. The southward Jupiter ap-proach results in a solar pass from the northalong an orbit that is 90° inclined from the eclip-tic plane and flies by the Sun at a speed of308 km/s, as shown in Figure 4-4. Its closest



Figure 4-4. Solar encounter trajectory details.

Figure 4-3. Mission timeline showing optional third solar encounterand relationships of the encounters to the solar cycle.

2010 2011 2012 2013 2014 2015 2016 2017 2018 2019 2020 2021

Predicted Solar Cycle

2010 2011 2012 2013 2014 2015 2016 2017 2018 2019 2020 2021 2022

LaunchMay-2010

7.1 years

JGAAug-2011

1st Sun FlybyJul-2013

2nd Sun FlybyJul-2017

TCM

3.1 years

¥ ¥ ¥3nd Sun Flyby

Jul-2021

11.1 years

2010 2011 2012 2013 2014 2015 2016 2017 2018 2019 2020 2021


2010 2011 2012 2013 2014 2015 2016 2017 2018 2019 2020 2021


2010 2011 2012 2013 2014 2015 2016 2017 2018 2019 2020 2021 2022

LaunchMay-2010

7.1 years

JGAAug-2011



TCM

3.1 years

¥ ¥ ¥3nd Sun Flyby

Jul-2021

11.1 years

2010 2011 2012 2013 2014 2015 2016 2017 2018 2019 2020 2021 2022

LaunchMay-2010

7.1 years

JGAAug-2011



TCM

3.1 years

¥• ¥• ¥•3nd Sun Flyby

Jul-2021

11.1 years

02-0817R-6

approach to the Sun is 3 RS from the surface near

the equator, and the pole-to-pole flyby takesplace within a solar distance of 9 R

S in a period

of 14 hours. The solar distances at 30 days and10 days before and after perihelion are 1 AUand 0.5 AU, respectively.

Both the solar encounter date and the orbit in-clination are selected to allow continuous real-time data transmissions to Earth during the solarpasses. As the spacecraft passes the Sun at clos-est approach, Earth is in quadrature, positioned90° from the Sun–probe line to the west of theSun. This configuration allows the spacecraft

passes occur every 4 years with the same favor-able quadrature geometry. A �V of 50 m/s isrequired for a burn performed 22 days after peri-helion, when the spacecraft is 0.8 AU from theSun.

�V Requirement. The onboard �V requirementis moderate. No deterministic �V is needed forachieving the first solar flyby, and only 50 m/sis required for targeting successive solar flybys.A �V budget of 150 m/s was estimated to cor-rect for launch vehicle dispersion, guidance, andnavigation errors. An additional 25 m/s was in-cluded for additional margin, resulting in a totalonboard capability of 225 m/s, excluding theseparately budgeted attitude control propellant.

4.2.3 Launch Vehicle Selection

The required C3 for the Solar Probe trajectory is

128 km2/s2. Two expendable launch vehicles(ELVs) can deliver this C

3 and satisfy regula-

tory constraints related to the use ofradioisotope thermoelectric generators (RTGs):the Atlas V 551 and the Delta IV 4040H-19. Thebaseline for this study is the Atlas V 551 withthe Star 48B third stage, although the design

simultaneously to point its an-tenna toward Earth and its ther-mal shield toward the Sun andit avoids direct impact fromdust particles that circulatenear the Sun in an east–westdirection. The perihelion timeis selected to have simulta-neous access to the Probe fromtwo Deep Space Mission Sys-tem (DSMS) stations, Gold-stone and Madrid (Figure 4-5).

A re-targeting burn will be per-formed after the first solarflyby to adjust the orbital pe-riod so that successive solar



Figure 4-5. DSMS coverage near perihelion.

0

10

20

30

40

50

60

70

80

90

-12 -10 -8 -6 -4 -2 0 2 4 6 8 10 12

Time from perihelion (hours)

Elevation (deg)

0

0.5

1

1.5

2

2.5

3

3.5

4

4.5SEP Angle (deg)

GoldstoneMadrid

Canberra

Sun-Earth-Probe Angle

0

10

20

30

40

50

60

70

80

90

-12 -10 -8 -6 -4 -2 0 2 4 6 8 10 12

Time from perihelion (hours)

Elevation (deg)

0

0.5

1

1.5

2

2.5

3

3.5

4

4.5SEP Angle (deg)

GoldstoneMadrid

Canberra

Sun-Earth -Probe Angle

02-0817R-8

could easily support the use of the Delta IV. TheAtlas V/Star 48B has an adequate predicted liftcapability of 713 kg, and its current advertisedcost is lower than that of the Delta. Figure 4-6shows the lift capability of the two launch ve-hicle options.

The spin-stabilized Star 48B third stage was se-lected over the 3-axis stabilized Star 48V largelybecause a spin-stabilized upper stage ispreferable from a range safety perspective forlaunches carrying nuclear payloads.

4.2.4 Mission Concept ofOperations

Launch and Early Operations.Solar Probe will be launchedfrom the Cape Canaveral AirForce Station within a 20-daylaunch window on a two-stageELV with a Star 48 third stage.Except for the propellant bud-geted for navigation correction,the entire �V necessary forachieving the Jupiter flyby willbe supplied by the launch ve-hicle. For the first- and second-

stage fly-out, the spacecraft will be 3-axis sta-bilized by the launch vehicle. For the third-stagefly-out, the stage will be spin-stabilized at ap-proximately 60 rpm.

Upon separation from the third stage, SolarProbe will perform a despin maneuver and as-sume a 3-axis stabilized orientation with themedium-gain antenna (MGA) and high-gain an-tenna (HGA) pointed toward Earth. The MGAwill be the primary antenna used for communi-cations for early checkout and cruise. The first90 days of the mission is allocated for early or-

bit checkout of the spacecraft,which can easily be supported by8-hour contact periods 3–4 timesper week.

Outbound Cruise. After the ini-tial checkout is completed, thenumber of required contacts isreduced to one 8-hour contactper week. The spacecraft re-mains 3-axis stabilized with theHGA and MGA pointed towardEarth. A slow spacecraft rotationwill be introduced about the an-tenna axis to minimize thecumulative angular momentumeffects of solar pressure torque.

0

200

400

600

800

1000

1200

1400

1600

1800

100 110 120 130 140 150 160 170

C3 (km2/s2)

Sep

arat

ed P

aylo

ad M

ass

(kg

)

Delta IV 4050H 19 with Star 48B

Atlas V 551 with Star 48B

Solar Probe Design Point (128, 713)

02-0817R-9

Figure 4-6. Solar Probe launch vehicle predicted lift capability.



Jupiter Gravity Assist. Beginning 30 days be-fore the Jupiter encounter, contacts will increaseto 3–4 per week to allow analysis, execution,and evaluation of a navigation burn planned for21 days prior to the flyby. This burn will pro-vide final corrections necessary to assure thatthe gravity assist will target the spacecraft forthe desired 4-R

S perihelion solar encounter.

Inbound Cruise. From 1 month after the Jupi-ter flyby until 3 months before the solar pass,the spacecraft returns to cruise mode and reducescontacts to one per week. At this time, contactsagain increase to 3–4 per week to perform nec-essary instrument calibrations and test spacecraftmodes planned for the encounter. At a distanceof approximately 0.8 AU, the spacecraft mustchange attitude to point the thermal protectionsystem (TPS) toward the Sun. Communicationwith Earth is maintained by a single-axis antennaarm and gimbal that controls the MGA and HGAposition and attitude. Freedom to point the an-tenna in the other axis will be attained by rotat-ing the spacecraft about the Z-axis (symmetricabout the thermal shield).

Solar Encounter. At 0.5 AU, 10 days prior toperihelion, the solar encounter phase begins, andground contact becomes nearly continuous. Thespacecraft will accommodate continuous datastreams from the instruments except during briefmomentum management maneuvers lasting lessthan 1 minute and no more frequent than every20 minutes. All science data are recorded re-dundantly on two solid-state recorders (SSRs).A portion of the science data is transmitted inreal time. During the encounter, the data rate willsupport at least 25 kbps. Spacecraft autonomywill maintain proper attitude for science collec-tion, adjust the HGA, and rotate the spacecraftto maintain communications with Earth. An ex-tensive autonomous fault protection system as-sures a safe spacecraft attitude in the event ofany single hardware or software failure.

Post Encounter. The solar encounter phase ends10 days after perihelion, and retrieval of non–real-time science data from the recorders begins.The spacecraft must maintain the TPS pointedtoward the Sun until 20 days after perihelion,when a retargeting burn of 50 m/s �V estab-lishes the 4-year orbital period and guaranteesquadrature for successive solar encounters. Thismaneuver will be performed while maintainingcontact with the ground.

After the retargeting burn, the spacecraft returnsto its cruise attitude configuration and resumesdownlinking the recorded science data.Downlinking the entire 128 Gbits of potentialdata will take approximately 60 days from theend of the encounter. The spacecraft remains incruise mode with weekly contacts until prepa-ration for the second solar encounter, which willbe operationally identical to the first encounter.Figure 4-7 summarizes operational concept as atimeline.

Cruise Mode Science. An additional capabilityof Solar Probe is collection of science data dur-ing the cruise phases. Since weekly contacts areplanned to support navigation and spacecrafthealth and maintenance, science data can becontinuously recorded and downlinked weekly.The spacecraft power and command and datahandling (C&DH) systems support instrumentoperation throughout cruise. Approximately 22Mbits of science data per week could betelemetered to the ground during these phases.

4.2.5 Mission Modes

Figure 4-8 shows the Solar Probe mission modesand mode transitions. These modes help definegroupings of spacecraft functionality useful foroperations planning, fault protection systemdevelopment, software development, and guid-ance and control (G&C).



Ascent Mode. While the spacecraft is in the fair-ing, all systems not required for initial deploy-ment are powered off and the RTGs operate at a

low power level. After initial fly-out, onboardtimers activate G&C components and thetransmitter. After deployment, the spacecraft

2010 2011 2012 2013 2014 2015 2016 2017 2018

LaunchMay 2010

JGAAug 2011

1st Sun FlybyJul 2013

2nd Sun FlybyJul 2017

¥ ¥

Early Orbit Checkout•3-4 contacts per week•~90 days

Jupiter Fly•Several contacts per week•~60 days

Data Retrieval•16-h coverage per day•~60 days to retrieve recorded data

Encounter 1 (P1± 10 d)• Nearly 24-h coverage

Cruise* Cruise* Cruise*

Encounter 2 (P2± 10 d)• Nearly 24-h coverage

Normal Ops.

Primary Science

Maneuvers

Encounter Prep.•Several contacts per wk•3 mos. prior to encounter

* Cruise periods assume one 8 hr contact per wee

Disp. Burn• ~ L+60 d

Navigation Burn• ~JGA – 3 wks

Retargeting Burn• P+20 d

2010 2011 2012 2013 2014 2015 2016 2017 2018

LaunchMa 2010

JGAAug 2011

1st Sun FlybyJul 2013

2nd Sun FlybyJul

¥• ¥•

•3-4 contacts per week•~90 days

Jupiter Flyby Prep•Several contacts per week•~60 days

Data Retrieval••

Encounter 1 (P1± 10 d)•

Cruise*Cruise* Cruise*Cruise* Cruise*Cruise*

Encounter 2 (P2± 10 d)•

Normal Ops.

Primary Science

Maneuvers

Encounter Prep.•Several contacts per wk•3 mos. prior to encounter

* Cruise periods assume one 8 h contact per week

Disp. Burn• ~ L+60 d

Navigation Burn• ~JGA – 3 wks

Retargeting Burn• P+20 d

02-0817R-10

Figure 4-7. Operational concept summary.

Figure 4-8. Solar Probe mission modes.

SCIENCE MODE

SPACECRAFT READY MODE

CRUISE SAFE MODE

SOLAR ENCOUNTERSAFE MODE

ASCENT MODE

MANEUVER MODE

Special Modes Normal Modes Fault Protection Modes

GroundCommand Ground

Command

Critical Fault

GroundCommand

Autonomous

Critical Fault

AutonomousFor Certain Faults

Autonomous

GroundCommand

02-0817R-13



autonomously performs a despin maneuver andpoints the antennas toward Earth. The spacecraftremains in Ascent Mode until a ground com-mand is received.

Spacecraft Ready Mode. The standard modeoutside of solar encounter periods is SpacecraftReady Mode, wherein the spacecraft has fullfunctionality and its attitude is normally 3-axisstabilized with the HGA/MGA pointed towardEarth.

Maneuver Mode. A ground command initiatestransition to Maneuver Mode, which is used toperform �V maneuvers both for navigation cor-rections and for the retargeting burn. Its special-ized functionality is isolated from the othermodes to prohibit accidental maneuvers. Tran-sition out of this mode is made autonomouslyafter completion of the maneuver sequence.

Science Mode. The normal mode during solarencounters and instrument calibrations is Sci-ence Mode, in which the instruments capturedata and are on the 1553 data bus schedule. Inthis mode, the TPS points toward the Sun andautonomous functions assure the safety of thespacecraft.

Cruise Safe Mode. One of the two safe modesfor Solar Probe, Cruise Safe Mode is used inthe event of a fault while the spacecraft is inSpacecraft Ready Mode. Nonessential loads arepowered off, and the spacecraft points its anten-nas toward the Sun using digital Sun sensors.Farther than 3 AU, this is sufficient to captureEarth within the MGA beamwidth. At closerdistances, a small step-stare search pattern willbe used to get the MGA to capture Earth. Atdistances less than 2 AU, a widebeam low-gainantenna (LGA) can also be used to reestablishcommunications. Cruise Safe Mode operates onsoftware isolated from the operational softwareand is triggered by the onboard fault protectionsystem.

Solar Encounter Safe Mode. This second safemode is only used during the solar encounters,and is triggered by fault protection logic activewhen the spacecraft is in Science Mode. Anindependent processor in the attitude interfaceunit (AIU) takes control of the G&C functionsusing a solar horizon sensor to maintain thespacecraft in a safe attitude. Depending on thefault, the primary systems or redundant systemswill take over, and promotion back to sciencemode will be autonomously performed. Forsome critical faults, ground intervention may berequired and mode transition would be from theground.

4.2.6 Summary

The Solar Probe mission design puts the space-craft in a favorable position to achieve the sci-ence objectives. The short cruise time fromlaunch to first solar encounter and the final or-bital period allows for two passes in 7 years,and possibly three coronal measurements withina single 11-year solar cycle. Quadrature geom-etry optimizes the downlink of science data dur-ing the encounters, making it unnecessary to relysolely on the transmission of recorded data af-ter the passes.

4.3 Spacecraft Overview

4.3.1 Spacecraft Concept Drivers

The Solar Probe spacecraft concept was con-strained by several significant design drivers anddefined many of the subsystem approaches.Most of the design drivers are based on the ex-treme environments the spacecraft will encoun-ter during the mission. Table 4-2 and theparagraphs below summarize the key designdrivers and reference the appropriate subsectionof this report where the resulting subsystem con-cept is discussed in more detail.

Solar Thermal Flux. The intense solar flux of400 W/cm2 at perihelion is the most challeng-ing spacecraft design driver. This heat input is



managed by a thermal protection system (TPS)that shadows the spacecraft bus and instrumentsfrom the direct solar flux. This TPS consists ofprimary and secondary thermal shields and lighttubes to allow nadir-pointing remote sensinginstruments to safely collect images while re-jecting additional flux that could damage them.

Solar Dust. Remote sensing instruments haveobserved dust near the Sun at latitudes of ± 30°from the equator with a generally a counterclock-wise orbit when observed from the Solar northpole. Based on remote observations of the solardust, an expected solar dust flux density modelwas developed based on the work of Mann(2001). Using this dust flux density model, thespacecraft was designed to withstand a singleimpact of a particle of 100-�m diameter per pass,plus additional impacts of smaller dust particles.Due to the relative velocities of the spacecraftand dust particles, impact velocities as high as400 km/s are likely. The design includes a massallocation for shielding of critical components,and the reaction wheels and thrusters are sizedto accommodate torques imposed by these im-pacts.

Solar Scintillation. The effects of solar scintil-lation have been well characterized based on

mission data provided by the NEAR-Shoemakerspacecraft during solar conjunction as well asearlier by Magellan and Galileo. During theNEAR mission, at solar conjunction, measur-able telemetry losses were experienced using X-band transmission once the angle between theSun, Earth, and the spacecraft came within 2.3°.For Solar Probe near perihelion, this angle willbe less than 2.3° for several hours, and anX-band downlink was not considered reliableto meet the real-time telemetry goal of 25 kbpsduring this most critical time in the mission. Ka-band transmission was selected as the baselinebecause it is relatively unaffected by the solarscintillation and could meet the 25 kbps goalduring the solar encounter.

Coronal Lighting. Expected coronal lightingnear the Sun is currently an important environ-mental uncertainty and can have significant con-sequences for maintaining attitude control.Excessive coronal lighting can increase back-ground noise and degrade the ability of the startracker to detect star constellations needed todetermine the spacecraft attitude. Coronal light-ing conditions can be estimated using data fromremote sensing instruments currently in orbit ata distance of 1 AU. This information may beenough to provide some characterization of the

Table 4-2. Solar Probe key design drivers and resulting concept impacts.

Requirement Resulting Environment Concept Impacts Subsection4 RS perihelion mission Solar flux ~ 400 W/cm2 • New heat shield development

• Light tubes for nadir imagers4.4.2

1% chance of ≥100-µm dustimpact

• Dust shield

Solar scintillation near peri-helion

• Ka-band downlink 4.4.3

Coronal lighting nearperihelion

• Alternative ACS safing sensor 4.4.7

93-krad total dose (100 mils)because of Jupiter flyby

• Radiation-tolerant electronics 4.4.5

≥2 solar encounters (>7.2-year mission life)

• Radioisotope power source• Redundant systems

4.4.4, 4.3.2

Pointing accuracy of 0.2°Inertial knowledge 0.05°Jitter of 0.005° in 1 s

• Star tracker• High-precision inertial measure-

ment unit (IMU)

4.4.7

Maximize science return • 128-Gbit recorder• Ka- and X-band downlink

4.4.3, 4.4.5

Mass loss rate ≤2.5 mg/s • 15° conical carbon–carbon heatshield

• Reaction wheels

4.4.2, 4.4.7



lighting environment; however, uncertainty willremain until a mission near the Sun is performed.To accommodate this uncertainty, the SolarProbe concept incorporates two star trackers,facing approximately orthogonal directions, anda high-precision IMU. Additionally, the conceptincorporates a safing sensor that detects the edgeof the solar horizon if the star trackers have beenblinded for an extensive period.

Radiation Environment. The Solar Probe ion-izing radiation environment consists of threecomponents: electrons and protons trapped inthe Jovian magnetosphere, solar protons duringsolar maximum conditions, and gamma raysfrom the RTG fuel. Table 4-3 provides the ion-izing dose behind 100 mils of Al spacecraftshielding using one-dimensional slab geometry.In this table, the Divine–Garrett model for a Jo-vian pass with minimum approach of 7 R

J was

used to calculate the radiation dose near Jupiter(Divine and Garret 1983). The JPL-91 solar pro-ton flux model (Xapsos et al. 1996) was used atthe 95% confidence level to estimate the dosecontribution for the period of the mission dur-ing solar maximum conditions. The RTG gammaray flux was derived from measurements of theCassini RTGs.

Both the relatively high total dose and the largeinitial dose at the Jupiter flyby affected the over-all concept. All electronics devices for this mis-sion are selected to be radiation tolerant to thetotal dose defined in Table 4-3. As a precaution,unnecessary radiation-sensitive components,including the redundant integrated electronics

module (IEM) and star tracker, will be turnedoff while the spacecraft passes through Jupiter’sradiation belts to minimize radiation effects.

Multiple Solar Encounters. The current mis-sion requirement is to conduct two solar encoun-ters. This design provides scientists theopportunity to collect data over different peri-ods of the solar cycle. Solar Probe incorporateshardware with an expected design life to accom-modate 2 or more encounters. In addition, thisdesign incorporates significant redundancy, dis-cussed in Section 4.3.2.4, to ensure mission suc-cess. Accommodating multiple solar encountersalso constrained the primary power source forthe mission. A radioisotope power source wasconsidered the only practical alternative to op-erate for more than 7 years in a widely varyingsolar flux environment of 0.02 to 5 AU.

Pointing and Jitter. The pointing accuracy re-quirement for Solar Probe was driven by twodifferent sources. The remote-sensing, solar-nadir-pointing instruments require an absolutepointing capability of 0.3° and attitude knowl-edge of 0.05°. Selection of Ka-band combinedwith the antenna dish size of 0.8 m resulted inan HGA pointing requirement of 0.2°. The AOalso specified a jitter requirement of maintain-ing boresight within 0.005° in 1 s during the solarencounter. The combination of these require-ments drove the selection of attitude determina-tion sensors to a star tracker and a high-precisionIMU. For attitude control, the selection of re-action wheels provides both very good attitudepointing control and low jitter.

Maximum Science Return. The principal goalin development of this Solar Probe concept wasto the maximize science return of the mission.Due to constraints in transmitter power and thesize of the HGA, real-time telemetry was lim-ited to 25–60 kbps and could not meet the sci-ence data needs alone. As an alternative,therefore, the concept incorporates redundant128-Gbit SSRs easily capable of duplicate re-cording of all the science data taken over the

Table 4-3. Solar Probe total radiation dose. The Ju-piter flyby radiation is the most significant componentof the total mission radiation dose.

Component Margin Dose with Margin(krad Si)

Jupiter pass (Divine–Garrett model)

×2 79

RTG gamma (scaledCassini data)

×1.5 4

Solar protons (JPL-91model)

×3 9.7

Total 93



entire solar encounter. With this concept, thefunction of the real-time telemetry is to providean important contingency capability and furtherincrease the potential science data return.

Maximum Spacecraft Mass Loss Rate. The1999 AO specified a maximum mass loss rateof the spacecraft materials of ≤2.5 mg/s duringthe solar encounter to ensure that spacecraftcharging and mass loading do not adversely af-fect the sensitive particle measurements beingtaken. At perihelion, this requirement is a sig-nificant driver; it limited the materials and maxi-mum temperature of the TPS, defining theprimary shield concept as a tall conical carbon–carbon structure. This requirement was also in-terpreted to include all spacecraft mass lossincluding propellant, which led to the selectionof reaction wheels over thruster attitude controlto minimize the propellant exhaust expelled into

the environment during sensitive particle sciencemeasurements.

4.3.2 System-Level Description

Solar Probe is a 3-axis stabilized spacecraft de-signed to operate in environments from 0.02 to5 AU from the Sun and in close proximity toJupiter (7 R

J). The concept is illustrated in Fig-

ure 4-9 in deployed configuration; spacecratdimensions are shown in Appendix A. The mostprominent feature is the TPS, with its large2.7-m diameter carbon-carbon conical primaryshield and a low conductivity secondary shieldattached at the base. This TPS protects the space-craft bus and instruments within its umbra dur-ing solar encounter. The TPS is attached to thebus via 12 struts. The spacecraft bus accommo-dates all instruments specified in the 1999 AO,including a retractable boom for several of the

Primary Heat Shield

Secondary Heat Shield

Thrust StrutsAxial Struts

High-Gain AntennaRTG (3)

Science Boom Launch Vehicle Adapter

02-0817R-41

Figure 4-9. Solar Probe external view deployed.



in situ instruments as well as the fail-safe solarhorizon sensor. A hexagonal/cylindrical bus ac-commodates the spacecraft subsystems and pro-vides an efficient mechanical structure to handlelaunch loads and integrate with the launch ve-hicle. The RTGs are placed high on the bus tohelp minimize the diameter of the TPS and arepositioned for integration with the spacecraft viaaccess doors in the fairing. The propellant tankis central to the bus, and most of the electronicsboxes are mounted on the bottom deck with thetelescoping boom.

The architecture for the Solar Probe concept(Figure 4-10) is derived from previous JHU/APLmissions, including NEAR, TIMED, STEREO,and MESSENGER. The dual IEMs are a cen-tral feature of these designs. They provide ahighly mass- and volume-efficient packagingsolution for the command and data handling

hardware, attitude control processor, andcommunications electronics. The instrumentsare grouped into remote sensing and in situsuites, each with a dedicated data processing unit(DPU) and power converter. Commands and dataare routed throughout the spacecraft through aredundant digital 1553 bus. Some attitude con-trol components require an analog interface. Theattitude interface unit (AIU) converts the ana-log and digital signals and provides the neces-sary interface to these components.

4.3.2.4 Fault Protection

Solar Probe fault protection was driven by sev-eral factors. First, considerable uncertainty re-mains in the solar encounter environmentsaffecting attitude control. Maintaining attitudecontrol and recovering rapidly from faults af-fecting the ACS are critical to spacecraft

Figure 4-10. Physical block diagram of the Solar Probe concept.

02-0817R-15

PCI Back plane

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VisibleMagnetograph

Helioseismograph

EUVImager

All-Sky 3-DCoronagraph

Magnetometer

Solar Wind IonComp & Electron

Spectrometer

Energetic ParticleCompositionSpectrometer

Plasma WaveSensor

Fast Solar WindIon Detector

Sun HorizonSensors

RemoteSensing

DPU(redundant)

PowerConverter

(redundant)

InSituDPU

(redundant)

PowerConverter

(redundant)

Fuse

Boa

rd

Rel

ay B

oard

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SRU(redundant)

RTGRTG

Li-IonBattery

RTGRTG

Power Distribution Unit (PDU)

(redundant)

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BoostConverter

ChargeController

Attitude Interface Unit (AIU)(redundant)

(redundant)

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survival. Second, critical aspects of the sciencedata collection occur over relatively short peri-ods. The flight through the coronal holes for eachencounter lasts only 14 hours. These consider-ations drive the spacecraft to having significantonboard autonomy and hot backup systems dur-ing the solar encounters.

The plan for two encounters, with a potential forthree or more, also drives hardware selection andredundancy considerations. Solar Probe incorpo-rates redundancy in virtually all subsystems. Theprimary single-point failure that remains is theRTG, for which it is prohibitively expensive tofly a spare unit. Even with a failed RTG, how-ever, a degraded mission is possible with the re-maining two units. All other spacecraft systemsare either redundant in their design or redundantunits are flown. Hardware redundancy is sum-marized in Table 4-4.

Years of space flight history have shown thatrandom failures constitute only a portion of mis-sion failures. Many anomalies have commoncauses or stem from design or manufacturingflaws; two identical, redundant units can bothfail. As a result, Solar Probe took more of a

“belt and suspenders” approach; the designincorporates many areas of functional redun-dancy that allow the spacecraft to either fullyor partially recover the mission using different(functionally redundant) hardware, even if bothredundant units fail. Solar Probe functional re-dundancy is highlighted in Table 4-5.

4.3.2.4 Resource Allocations

Mass Allocations. Table 4-6 shows the masssummary for Solar Probe. Mass is summarizedfor the instruments and each major subsystemand is compared against the maximum launchmass of 713 kg. The mass estimate shows thecurrent estimate of the mass at launch. Thegrowth allowance represents the allowable per-centage growth of that subsystem based on thematurity of components within that subsystem.The instruments were provided a generous massgrowth allowance of 40%. Generally, new ormodified hardware on the spacecraft was givena 20% growth allowance, and components thatwould be procurements of existing hardwarewere given 10% growth allowance. The not-to-exceed (NTE) mass represents the mass with thegrowth allowances incorporated.

Table 4-4. Hardware redundancy.

Functional Area Hardware Redundancy CommentsC&DH processing 2 processors • Hot spare during solar encounter

• G&C processing on same unitG&C processing 2 full-function processors

2 safing processors• Hot spare during solar encounter• C&DH processing on same unit• Different safing processors and algorithms

Attitude determination (cruise) 2 star trackersSun sensors

• 1 star tracker turned off during Jupiterflyby

Attitude determination (solarencounter)

2 star trackersSolar horizon sensor

• Both star trackers on• Solar horizon sensor always on

Attitude control 4 reaction wheels • Three needed for 3-axis controlPropulsion 16 thrusters • Redundant coupled pair in each axisData bus Redundant 1553 busData storage 2 SSRs • 1 recorder turned off during Jupiter flyby

• Science data duplicated on each recorderduring solar encounter

Telecommunications 2 uplink cards2 downlink cards2 LGAsMGA & HGA2 X-band solid-state power ampli-

fiers (SSPAs)2 Ka-band SSPAs

• Both receivers on at all times• Either MGA or HGA can be used for

cruise communications• Redundancy in X- or Ka-band to either

MGA or HGA



individual components, is given in the masterequipment list (Appendix A).

Power Summary. Loads imposed during the so-lar encounters drive the spacecraft power requiredfor the mission. Basically, during this part of themission, all subsystems and instruments are ac-tive. In addition, because of the extreme solarpressures and solar dust impacts, frequent mo-mentum management maneuvers will be needed,requiring additional power for the thrusters.

Table 4-7 summarizes the power required dur-ing the solar encounter and compares it with thepower available with the three RTGs for eachsolar encounter. As shown, significant marginin average power is available for both solar en-counters. Occasionally, transient peak loads,mostly from occasional thruster firings and re-action wheel torques, will exceed the power pro-vided by the RTGs alone. A 5 A-h batteryincorporated as secondary power source easilyaccommodates these transient loads. A more de-tailed breakdown of power usage during themission is provided in Appendix A.

Table 4-5. Added functional redundancy in Solar Probe concept.

Functional Area Primary SystemFailure Functional Redundancy Mission Impact

G&C processing G&C softwarefault

Different AIU processor withindependent G&C software

• Reduced pointing capabilityuntil new software loaded

• Unable to execute ∆Vmaneuvers

Star tracker IMU – temporarySun sensors

• Maintaining communicationswith ground more complex

Attitude determination(cruise)

IMU Star trackerSun sensors

• No significant impact

Star tracker IMU – temporarySolar horizon sensor

• Eventual loss of HGA pointingcapability

Attitude determination (so-lar encounter)

IMU Star tracker • Reduced pointing capabilityAttitude control Reaction wheels Thrusters • Increased propellant usage

• Sublimation requirement maynot be met

Data storage SSRs Real-time telemetry • Reduced science dataKa-band downlink X-band downlink • Potential real-time data loss

near perihelionX-band downlink Ka-band downlink • No significant impact

Telecommunications

HGA MGA • Reduced real-time data duringsolar encounter

• Longer data retrieval timesPower Battery hardware RTG power alone • Increased probability of need

to power cycle instruments• Unnecessary hardware turned

off during maneuvers

Table 4-6. Solar Probe mass summary.

Subsystem Mass (kg)

GrowthAllowance

NTE Mass(kg)

Instruments 39.70 40% 55.58Telecomm. 16.92 20% 20.30G&C 40.58 19% 48.24Power 95.80 20% 114.96Structure 85.60 20% 102.72Thermal control 148.80 28% 190.96C&DH 14.41 20% 17.30Propulsion 21.83 11% 24.42Harness 21.55 20% 25.86Dust shield 12.00 20% 14.40Total dry mass 497.19 24% 614.73Maneuver propellant 63.10 63.10ACS propellant 6.00 6.00Trapped propellant 1.37 1.37GN2 0.60 0.60Total expendables 71.07 71.07Total wet 568.26 685.80

Max launch 713.00Reserves 27.20Total dry mass growth margin 29%

Propellant masses are estimated based on therequired �V and ACS propellant assuming themaximum launch mass of 713 kg. Mission re-serves represent unallocated mass that can beapplied to the overall system growth margin. Amore complete listing, with mass breakdown of



coronagraph imager (ASCI) are located on the+X side of the spacecraft, allowing them to pointin the velocity direction. The energetic particlecomposition spectrometer (EPCS), visible mag-netograph–helioseismograph (VMH), and ex-treme ultraviolet imager (EUVI) are locatedtowards the –X side to help balance the configu-ration. The design incorporates a retractableboom that accommodates the magnetometer,plasma wave sensor (PWS), and fast solar windion detector (FSWID). The instrument boom iscapable of extending from 1.5 to 5 m relative tothe bottom of the spacecraft bus. The boom willextend and retract during the solar encountersto match the maximum bus standoff distanceavailable within the protective umbra of the TPS.All instruments are located with a clear FOV asdefined in the AO (Figure 4-12).

4.4 Spacecraft Subsystem Descriptions

4.4.1 Mechanical Description

4.4.1.1 Requirements

In the development of the Solar Probe concept,several requirements constrained and drove themechanical configuration. In addition to themechanically related mission requirements de-scribed in Section 4.1 (Table 4-1, requirements7–13), several requirements were derived to sup-port the selected mission implementation andare included in Table 4-8.

4.4.1.2 Instrument Accommodation

Instrument mounting locations are shown in Fig-ure 4-11. The solar wind ion composition andelectron spectrometer (SWICES) and all-sky

Table 4-8. Mechanical subsystem requirements summary.

Requirement Mechanical ImplementationAccommodate instruments as defined in AO • Mount in locations with clear FOVs

• Provide retractable boom for in situ instrumentsPackage instruments and subsystems within TPSprotective umbra

• Compact design (0.95 m bus diameter)• RTGs mounted high under TPS

Provide single-axis HGA range of motion of 0° to 33° offof the nominal –Y boresight and maintain it withinprotective umbra

• Single-axis gimbaled HGA with two joints and exten-sion arm

Meet interface requirements of Atlas V or Delta IV • Spacecraft fits well within 5-m fairing dimensions• STAR 48B stage integration straightforward• Low center of gravity offsets

Support TPS, propulsion RTG launch loads, and transferinto launch vehicle adapter

• Thrust struts provided to support TPS• Brackets designed to support RTG and propulsion tank

Table 4-7. Power summary during solar encounter.

Subsystem Average Estimated PowerDuring Solar Encounter (W)

GrowthAllowance (%)

NTE Average Power Dur-ing Solar Encounter (W)

Transient PeakLoads (W)

Instruments 39.2 20 47.0 100.0Power 25.6 17 30.0 30.0Guidance & control 66.8 10 73.5 191.6Propulsion 9.5 10 10.5 46.8Telecommunications 26.1 18 30.8 30.8C&DH 47.8 10 52.6 52.6Harness losses 3.2 14 3.7 6.8Total average power 218.2 14 248.0 458.6

Power available 1st encounter (W) 293.7 Power reserves 1st encounter (W) 45.7 –164.9

Power growth margin 1st encounter 35%

Power available 2nd encounter (W) 273.9 Power reserves 2nd encounter (W) 25.9 –184.7

Power growth margin 2nd encounter 26%



SWICES Plasma Deflector(extending outside of umbra)

SWICES ObliqueSpectrometer

SWICES Nadir Spectrometer

ASCIEPCS

EUVI Detectorand light tube

VMH Detectorand light tube

FSWID Sensor

PWS Orthogonal Search Coils

PWS Orthogonal Antennas

Magnetometer

In-Situ and Remote sensing DPU sand Power converter electronicsinside of primary structure

Telescoping science boom(1.5 m at perihelion)

(Solar Horizon Sensor on the endof the boom)

02-0817R-16

Figure 4-12. Instrument fields of view.

Figure 4-11. Instrument mechanical accommodations.

EUVI Detector NadirPointing through 3 cmlight tube, FOV 0.6°

VMH Magnetograph Nadirpointing through 5 cmlight tube, FOV 0.3°

EPCS orthogonalto Z axis with 135°X 300° FOV

PWS monopole and dualdipole antennas on outerboom surfaceFSWID 4� sr FOV

ASCI Ram pointedHemispherical FOV(With partial view ofSWICES and heat shield)

SWICES Nadirpointing FOV –20°(With partial view of15° shield)

SWICES ObliqueSpectrometer Rampointed 135° X 300° FOV

Star Trackers Orthogonalto primary axis with clearFOV away from Ram

20°

+Z

+X RAM

02-0817R-46



4.4.1.3 Subsystem Accommodation

Major spacecraft subsystem locations are shownin Figure 4-13. Most of the subsystem electron-ics fit within the interior of the hexagon panelstructure. The RTGs are mounted externally toprovide adequate view of radiator surfaces tospace for cooling. The RTGs also need to beaccessible for installation through fairing doorswhile on the launch pad. The New Horizonsmission to Pluto is using a similar arrangementfor its single RTG and has shown that this typeof mounting is strictly required to do installa-tion on the pad through the fairing doors. Majorcomponents of the propulsion system includethe Centaur diaphragm tank and four rocket en-gine modules (REMs). Plumbing is routed in-side the structure and allows removal of sidepanels. The structure can also be integrally as-sembled with the propulsion system so stainless

steel tubing runs can be welded in place fromthe tanks to the thrusters.

All subsystems fit within a 16.5° umbra providedby the TPS at perihelion. Nominal umbra at peri-helion is 14.5° with 2° allocated to account formisalignments and attitude control errors. Thediameter of the primary heat shield (2.7 m) isdriven by the need to provide umbra shadow overthe RTGs.

Antenna placements are defined by mission ori-entation, with the HGA, MGA, and one LGApointing along the –Y-axis toward Earth. AnotherLGA is on the opposite side to supply sphericalantenna coverage for tumble recovery. Agimbaled joint for the HGA and MGA assemblyprovides single-axis rotation in the Y-Z plane forEarth tracking. A second rotation joint allows theHGA to move out from the bus, allowing the full

Figure 4-13. Spacecraft subsystem mechanical accommodations.

TPS Axial SupportStruts (6)

Telescoping RetractableScience Boom (1.5-4 m)

Single Propellant Tank(Centaur Tank)

TPS Carbon-Carbon Primary Heat Shield; 2.7 m diameter

TPS Secondary Shield19-cm Aerogel SandwichedBetween Carbon-Carbon Face Sheets

0.8-m High-Gain AntennaSingle Axis Gimbal Arrangement On Boom (0¡ to 60¡ Range)

All components contained within umbraby 1.5° at 4 RS for nominal solar nadirpointing

MMRTG (3)

Star Tracker (2)

94 cm (37-in.) PayloadAdapter Ring

Reaction Wheel (4)

Avionics, Power DistributionBoxes on Panel Tucked IntoCylinder

Medium-Gain AntennaGimbaled with HGA

Low-Gain Antenna (2)

Solar Limb Detector

Rocket Engine Module (8)

TPS Thrust Struts (6)

02-0817R-42



0° to 33° of rotational freedom necessary to main-tain contact throughout the solar encounter whilekeeping the heat shield pointed toward solarnadir. Drive components for the antenna systemuse off-the-shelf stepper motors that have veryaccurate pointing capability, are highly reliable,and have heritage on numerous programs. Alaunch restraint mechanism attaches the HGA tothe structure and will be commanded for a one-time release on orbit.

4.4.1.4 Bus Mechanical and StructuralDesign

The primary structure for the spacecraft bus useslow-mass aluminum honeycomb panels andmachined aluminum construction typical of re-cent JHU/APL spacecraft. Detailed structure de-sign methodologies, materials, and attachmentsdraw heavily from past and current programsand take advantage of new technology that mini-mizes mass while maintaining low technical,cost, and schedule risk. Panels are designed foreasy removal to facilitate access to subsystemand instruments inside the structure. The panelsalso provide good thermal conduction andelectromagnetic compatibility (EMC) ground-ing for electronics and instruments.

The primary structure also has very efficient loadpaths to carry subsystems, instruments, and theTPS. The two main structural components are thecylindrical launch adapter and the hexagonalequipment module. This structure is sized to carrythe launch loads and meet launch vehicle stiff-ness requirements. The single Centaur propulsiontank mounts to the base of the hexagon withbrackets and efficiently transfers load directly intothe adapter ring. The launch vehicle adapter ringis a one-piece aluminum 7075-T73 machined ringforging typical of most launch adapters and pro-vides a very high strength factor of safety. Thehexagon consists of structural panels with 1.0 to1.5 mm aluminum face sheets bonded to light-weight aluminum honeycomb core. Interconnec-tion of panels is made with aluminum clips that

are integrally bonded to the panels. This designhas tremendous flexibility and can be modifiedlate in the schedule to accommodate changes ininstrument and subsystem interfaces.

Loads from the TPS are efficiently transferredfrom the strut system into the launch vehicleadapter. After achieving orbit, the thrust strutsare baselined to be thermally decoupled fromthe spacecraft body to limit heat transfer fromthe heat shield into the bus during solar encoun-ters. Options include simply disconnecting thestrut from the bus or ejecting the strut altogether.The latter approach may be advantageous be-cause it may reduce possible FOV or thrusterimpingements. A more detailed design of theTPS support system is needed, along with fur-ther FOV study, to determine which approach(if any) is needed to decouple the thrust strutsfrom the bus after launch.

4.4.1.5 Thermal Protection SystemStructural Analysis

Initial structural analysis was performed to aidin the design of the TPS and its interface to thespacecraft bus. This analysis was the structuraldriver in meeting both Atlas V and Delta IVlaunch environments.

The quasi-static limit load factors, sinusoidalvibration environment, flight acoustics, payloadseparation shock, ascent pressure profile, andfrequency requirements of both vehicles wereconsidered. Positive margins of safety will bedemonstrated for all load conditions and com-binations, and appropriate factors of safety willbe designed into all structural components. Afinite-element model of the primary shield, sec-ondary shield, and struts was developed inNASTRAN.

The primary shield consists of six plies ofcarbon–carbon in a two-dimensional quasi-iso-tropic layup (0, +60, –60) with a thickness of0.8 mm. An internal ring stiffener was added atthe middle of the conical shield to alleviate



4.4.2 Thermal Design

4.4.2.1 Requirements and Drivers

The Solar Probe thermal design presents aunique thermal design challenge set by the ex-treme environmental requirements (Table 4-1,requirements 1, 17, and 18). The spacecraft com-ponents and instruments must be kept between–20° and +50°C, the propulsion system between10° and 40°C. These temperatures must bemaintained over extreme solar flux environ-ments from 4 R

S to 5 AU. In addition, the ther-

mal system needs to reject excess heat providedby three multi-mission RTGs mounted to the busstructure, on average 248 W of internal thermaldissipation, and a scoop connected with the so-lar wind ion composition and electron spectrom-eter that is exposed directly to the solar flux.

An important design constraint in developing athermal protection system was to limit the out-gassing of all the thermal components to ≤2.5mg/s. This can be a very limiting constraintwhen developing thermal shields to operate inmaximum solar fluxes of 400 W/cm2.

4.4.2.2 Thermal Protection System (TPS)

The spacecraft must be shielded from solarheating during the near-Sun mission phases.The thermal control design approach incorpo-rates a TPS that shields the spacecraft from theSun at distances closer than 0.8 AU. At 4 R

S,

the solar radiation is 400 W/cm2, 3000 timesstronger than it is at the Earth. The TPS con-sists of a primary shield, secondary shield, andlight tubes, and it is integrated to the space-craft by 12 struts. The primary shield consistsof a high-aspect-ratio cone that keeps the peaksystem temperature below 2200 K. The sec-ondary shield provides the bulk of the thermalinsulation between the primary shield andspacecraft bus. The light tubes limit incidentflux to the two nadir-pointing imaging instru-ments. The struts provide support and stiffness

local modes in the primary shield while mini-mizing added mass. This stiffener is a 1 � 1 in.square closed section, with a wall thickness of0.5 mm. The material is a two-dimensionalwarp-aligned 5:1 carbon–carbon that achievesa high elastic modulus in the longitudinal direc-tion and is the same material being used for thestruts. The 12-strut arrangement contains sixstruts for thrust loads and six for lateral loads.Each strut is 5 cm in diameter and 0.5 mm thick,a geometry that balances thermal and structuralrequirements. The total mass of the shield ar-rangement is 121 kg.

The TPS was designed to withstand thrust loadsof 20 G and lateral loads of 12 G, but it is pri-marily a stiffness-driven design. Secondarystructure mode frequencies must be above 35Hz to avoid undesirable coupling with launchvehicle modes and/or large fairing-to-spacecraftrelative dynamic deflections. The NASTRANfinite element model of the heat shield assem-bly attached to the spacecraft bus calculates aprimary structural mode of 49.8 Hz, demonstrat-ing compliance with this requirement.

4.4.1.6 Spacecraft Handling

Handling of Solar Probe during test requires theTPS to be removed from the spacecraft for ship-ment to the launch site. JHU/APL has existingcontainers and fixtures from previous missionsthat would accommodate both the TPS and thebus to meet the handling requirements. A cus-tom lift fixture for the TPS will attach at the mid-point stiffening ring and allow easy integrationonto the spacecraft.

4.4.1.7 Summary

The mechanical concept meets all mission andderived requirements and represents a robustdesign with an overall low technical risk. Addi-tional definition of the mechanical design willrequire interaction with the instrument teams.



during launch and during the mission. Half ofthe struts, which are included to support the TPSonly for launch loads, will be thermallydecoupled after launch.

The TPS concept is shown in Figure 4-14. Inthis depiction, six of the struts are not shown,since they will be thermally decoupled afterlaunch, and the light tubes are integrated withinthe primary shield and secondary shields.

The primary shield is a 15° half-cone that is al-ways pointed toward the Sun when the spacecraftis closer than 0.8 AU and is constructed com-pletely of carbon-carbon composites. It is 2.7 min diameter with a 0.8-mm skin thickness. Asnoted above, a primary mission science require-ment is to have a minimal effect on the localplasma density. This is accomplished with a 15°half-angle cone, which has a small area projected

to the Sun but a large total surface area. The ef-fect of cone half angle on primary heat shieldtemperature is shown in Figure 4-15. The result-ing TPS carbon-carbon outgassing rate as a func-tion of cone angle is shown in Figure 4-16. Massloss rates are below the required value of 2.5 mg/s for cone half angles up to 22°. However, theuncertainty in the design values and the increas-ing sensitivity with temperature argue for a con-servative approach. Therefore, the 15° cone halfangle was chosen for the design study.

The objective of the secondary shield is to limitthe heat flow between the primary shield and thespacecraft bus; also, separating the secondaryshield from the spacecraft bus allowsthe thermal energy to radiate into space. Thesecondary shield consists of a thick, low-density, low-conductivity material contained bycarbon-carbon face sheets that provide support.It closes off the base of the cone formed by theprimary shield. The current concept calls for a20-cm thickness. The hot side shield tempera-ture reaches the same level as the primary shield,but the cold side drops to between 500 and600 K. The secondary shield is separated fromthe spacecraft bus by 0.5 m. The remainingspacecraft insulation consists of the high-tem-perature multilayer insulation (MLI) on the topsurface of the bus.

Two material choices are available for the sec-ondary shield. Both carbon–carbon aerogel andspun carbon fiber batts represent modificationsof existing technologies requiring some devel-opment (Frazer 2002). A critical feature of theTPS is that the planned development work con-tinue on both materials. The design values forthe secondary shield density and conductivityassume 0.06 g/cm2 and 0.1 W/m K, respectively.

The aerogel technology is currently being devel-oped by Southern Research Institute (SRI). SRIefforts have been conducted using in-house fund-ing and are limited in temperature. Present con-ductivity measurements support values as low asFigure 4-14. TPS mounted to spacecraft bus.

02-0817R-17



0.04 W/m K. Of concern at higher temperaturesare the microstructural changes that affect boththe conductivity and density. Studies indicate thatthese changes are relatively constant to 2300 K.As part of the risk mitigation effort, SRI will con-tinue their work in high-temperature aerogels,

which will include material reformulation andfabrication studies and high-temperature mate-rial property testing.

The other possible choice of material for thesecondary shield is carbon fiber batts. The

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Figure 4-15. Effect of cone half angle on primary heat shield temperature.

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02-0817R-19

Figure 4-16. Effect of cone half angle on carbon mass loss rate.



Applied Science Laboratory has proposed todevelop this material. Spun fibers of approxi-mately 1-�m diameter are fabricated into ma-terials with bulk densities of approximately0.05 g/cm2. The fiber orientation and densitymean the bulk of the heat transfer will be byradiation. Bulk conductivity values arecalculated from radiation theory at about0.02 W/m K. The analysis is supported withtesting on similar material systems. Risk miti-gation efforts include fabrication of thefibers into 20-cm bats and material propertytesting after exposure to 1 G and vibrationenvironments.

Two carbon–carbon light tubes are integrated atthe top to the primary shield and run through thesecondary shield, stopping at the bottom carbon-carbon face sheet. Maintaining a gap between theinstruments and the light tubes is necessary tolimit conductive thermal flux into the bus.

4.4.2.3 Thermal Modeling and Analysis

The baseline thermal design was modeled andanalyzed (Giacobbe 2002) for both the hot caseat 4 R

S from the Sun and at 5 AU during aph-

elion. All the analyses assume steady-state con-ditions at the given Sun distance. Theabsorptivity-to-emissivity ratio (�/�) of the pri-mary shield was assumed to be 1 and representsan uncoated carbon–carbon surface. The con-ductivity of the carbon–carbon was modeled asa function of temperature. There are six supportstruts. Each strut is 1.7 m long, with a 5-cmdiameter and a 0.5-mm wall thickness. The busstructure is primarily covered with MLI thermalblankets with 1.13 m2 surface area exposed asradiators.

The thermal analysis of the light tubes (Lee andLee 2002) shows that the spacecraft heat inputcomes both from direct solar radiation down thetube and re-radiation from the sides of the tube.The direct insolation is about 23 W. The sec-ondary radiation from the tube ranges from 1 to

25 W depending on the length of the tube. Thereis less re-radiation at larger spacings betweenthe light tube end and the spacecraft deck. How-ever, optical issues also drive baffle design andwill affect light tube analysis. A 75-W light tubeheat input was assumed for the study.

A thermal scoop associated with the solar windion composition and electron spectrometer isalso included to deflect local particles towardthe instrument apertures. For this study, we as-sumed that the instrument team would designthe scoop and be responsible for limiting ther-mal input into the bus. However, the scoop canhave a significant thermal impact on the space-craft thermal balance. The scoop is also madeof carbon–carbon, with the exposed portionreaching the same temperatures seen by the pri-mary shield. Without thermal shielding, the heatinput from the scoop will raise the spacecrafttemperature by several hundred degrees C.Analysis indicates the heat input can be reducedto acceptable levels with a 10-cm-thick, hemi-spherical, aerogel shield. Thermal isolation atthe mounting interface remains a critical issue.Because the instrument scoop is critical and be-cause the test efforts for the scoop and the pri-mary shield are similar, we recommend that thespacecraft design team take a larger role in scoopdesign and fabrication.

The spacecraft bus structure consists of a hex-agonal honeycomb panel structure mounted tothe cylindrical launch vehicle adapter. Theshadow of the TPS protects all spacecraft andspacecraft-mounted components from direct so-lar input. The RTGs are mounted to three of thesides of the hexagonal structure and are ther-mally isolated. The spacecraft support electron-ics are mounted near the bottom of the hexagonalstructure and in the adapter section. Electricalpower not required by the spacecraft is rejecteddirectly to space with two shunt panels that arethermally isolated from the spacecraft. The pro-pulsion section is thermally tied to the space-



craft bus to eliminate the need for separate heat-ers on the tanks, lines, and thrusters. While re-ducing heater power, this configuration requiresthe bus temperature to be kept above 10°C. Al-though the design includes the option of using avariable-conductance heat pipe/radiator systemto reduce the heat leak during the cold part ofthe mission, the current analysis does not includethis feature because the design is feasible with-out it. An option for a variable-conductance heat

pipe system represents additional margin in thespacecraft design.

Key features of the thermal design and the ther-mal balance for the hot case are illustrated inFigure 4-17. Although over 3400 W is trans-ferred from the primary shield to the secondaryshield, most of this heat is radiated to space. Lessthan 25 W is transferred from the TPS to thespacecraft bus, and most of this comes from the

Figure 4-17. Thermal balance for the hot analysis case.

02-0817R-20



02-0817R-21

TPS support struts. The spacecraft is effectivelyisolated from the primary shield by the second-ary shield, the low-� surfaces on the secondaryshield, and the spacecraft MLI. Internal dissi-pation provides the bulk of the heat input to thespacecraft, and the RTG provides the second-largest heat input. The RTG is assumed to bethermally isolated, but the actual heat flow canbe tailored based on the conductivity of themount and the needs of the system.

Hot case temperature analysis results are illus-trated in Figure 4-18 and shown in Table 4-9.Results for the baseline concepts indicate the

spacecraft bus temperatures will be maintainedwithin the required temperature limits.

Cold case temperature analysis results are illus-trated in Figure 4-19 and shown in Table 4-9.The results show acceptable temperatures at5 AU. Spacecraft bus temperatures ranges from–16 to 17°C, with the propulsion tanks 5°C. Ifrequired by the final spacecraft design, the pro-pulsion system temperatures could be raised withthe use of a variable-conductance heat pipe tolimit the heat loss from the spacecraft in the coldcase. The cold case temperatures include thereplacement heater power to keep the bus

Figure 4-18. Hot case thermal analysis results.



02-0817R22

propulsion system components to raise their lo-cal temperature during the cruise phase.

4.4.2.4 Summary

The thermal concept defined meets the missionrequirements with reasonable technology risks.An advantage of the MMRTG power source isthe ability to use the waste heat to help maintainspacecraft bus temperatures at large solar dis-tances without the need for additional heaters.There are identified technology risks in theimplementation of the thermal protection sys-tem. Section 5, Risk Mitigation, discusses thenext technical development activities, which will

Figure 4-19. Cold case thermal analysis results.

dissipation at a near-constant 248 W.Heater power in the cold case is moved to the

Table 4-9. Spacecraft temperature results ofthermal analysis.

Spacecraft Location Hot CaseTemp. (°C)

Cold CaseTemp. (°C)

Primary shield 1879 –167Secondary shield 440 to1879 –160Struts 237 –112Top deck MLI 88 –36Top deck 53 –13Hex structure 16 to 37 –17 to 8Adapter structure 8 to 27 –16 to 17Bottom deck 9 to 41 –13 to 3Propellant tank 42 5 to 6HGA –18 –35



reduce risk by fabricating and demonstrating aprototype design early in development.

4.4.3 Telecommunications Subsystem

4.4.3.1 Requirements and Constraints

Several requirements and constraints drove theSolar Probe telecommunications design. Theprimary drivers and resulting implementation aresummarized in Table 4-10.

4.4.3.1. Design Considerations

Downlink Frequency Considerations. One ofthe most important considerations in selectingthe telecommunications architecture was theneed to provide real-time telemetry near peri-helion. As Solar Probe approaches perihelionduring a solar encounter, the Earth is in quadra-ture with the Probe–Sun line, and the elonga-tion (Sun-Earth-Probe or SEP) angle isapproximately 1°. Conjunction experiments withother probes show that X-band and S-band com-munication links are affected adversely at smallelongation angles. For example, a study of theNEAR-Shoemaker spacecraft as it underwent asuperior conjunction in early 1997 showed mea-surable degradation of command and telemetrylinks at X-band for elongation angles below 2.3°(Bokulic and Moore 1999). On the downlink and

at an elongation angle of 1.1°, the ground sta-tion successfully recovered only 3% of the te-lemetry frames transmitted by the spacecraft at1104 bps, and 0% of the same when transmittedat 39.4 bps. The predicted telemetry marginswere 6 dB at 1104 bps and 19 dB at 39.4 bps ifsolar effects were ignored. For elongation anglesgreater than 2.3°, the ground station recoveredin excess of 99.7% of the transmitted frames at1104 bps. The published literature characterizessuch degradation as a scintillation loss. Earlierstudies on the Magellan (Webster 1994) andGalileo (Beyer et al. 1996) spacecraft producedsimilar conclusions.

A model of radio wave propagation in turbulentmedia (Armstrong and Woo 1980; Koerner1984) generates similar conclusions about fre-quency dependence of such degradation and pre-dicts that Ka-band links will be significantlymore resistant to corona scintillation effects thanX-band and S-band links. Improvements in linkperformance at Ka-band have been confirmedby simultaneous transmission of telemetry on X-band and Ka-band links during solar conjunc-tions with Mars Global Surveyor (MGS)(Morabito et al. 2000); Deep Space 1 (DS-1)(Morabito et al. 2001); and Cassini (Morabito etal. 2002).

According to models developed from the data(Armstrong and Woo 1980), scintillation lossvaries inversely as the square of the frequency.The frequency ratio of a Ka-band frequency toX-band frequency for a given DSMS channel is3.8:1, and consequently, the model predicts thatKa-band enjoys a 11.6-dB advantage overX-band with respect to scintillation loss. Poorweather does erode some of the advantages ofKa-band, but even under 90% weather condi-tions, only a Ka-band system can meet real-timetelemetry goals near perihelion. For this reason,the Solar Probe concept is baselined to use a Ka-band downlink during critical periods of solarencounters.

Table 4-10. Telecommunication subsystemrequirements summary.

Requirement ImpactMaximize science return overthe mission

128 Gbitbaselined

Maximize real-time telemetrythroughout a solar encounter

≥ 25 kbpsbaselined

Solar corona effects (scintilla-tion) on radio wave propagation

Ka-banddata down-

linkExtreme speed of the space-craft

X-bandcommand

uplinkLimited size of TPS to shadowthe antennas

0.8-m an-tenna di-ameter

Tight power constraints associ-ated with an RTG-powered mis-sion

8-W Ka-band

transmitter



An X-band system has advantages when solarscintillation is absent. An X-band system, whentaking advantage of the 70-m DSMS antennas,will actually outperform a Ka-band system,which is limited to the 34-m DSMS antennasfor the foreseeable future. In addition, X-bandis relatively insensitive to weather as comparedwith Ka-band, offering additional flexibility. X-band also offers the benefit of greater techno-logical maturity and flight heritage. And finally,an X-band system has a wider beam for a givenantenna size than Ka-band, making pointingbudgets more forgiving. For these reasons, it wasconsidered advantageous to carry an X-banddownlink in the baseline design, as well as a Ka-band downlink.

Figure 4-20 shows the performance of thebaseline telecommunications concept during asingle-day solar encounter. The graph assumesan 8-W RF transmitter for X- and Ka-band. TheX-band transmission is received using the 70-mantenna, and the Ka-band transmission is re-ceived using the 34-m antenna. Both are shownfor clear weather conditions (cumulative distri-bution [CD] = 0.1, 0.9), and we further showKa-band at 90% condition. In all cases, the cal-culated data rates were achieved with a link

Figure 4-20. Comparison of Ka-band and X-band links during a solarencounter.

margin of 3 dB. Scintillation losses and varyingelevations of ground track stations are includedin the calculations.

Over the range of weather (CD = 10% to 90%),the Ka-band link comfortably meets the mini-mum data rate over the entire primary scienceperiod, albeit with slight deterioration of linkmargin from 3 to 2.45 dB during the early partof the solar encounter. Figure 4-20 further showsthat the X-band link outperforms the Ka-bandlink except during the 24-hour (critical science)period around perihelion. Outside of this period,the data rate exceeds 45 kbps and is less suscep-tible to variations in weather and elevation angle.

To reduce the sensitivity of link performance tomonth of arrival, our architecture utilizes sci-ence telemetry links at both X- and Ka-bands.Provision for both frequencies using the sameantenna aperture comes at the expense of some-what reduced performance for X-band. Never-theless, the reduced X-band data rates outsidethe single day around perihelion remain com-parable to those of the Ka-band link but are lesssensitive to weather.

Command Link Considerations. The desire for

Data Rate at Ka- & X-band around perihelion

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Attribute X-band Ka-bandAntenna 70 m 34 mTX loss 1.5 dB 2.0 dBPointing 0.08 dB 1.1 dB

02-0817R-23

simplicity and the extremespeed of the spacecraft at peri-helion drive the choice of up-link frequency. The spacecraftreaches a top speed of 308 km/s, at which the angular separa-tion between transmitting andreceiving line of sight is nearly3.5 beamwidths of a 34-m an-tenna if both links are at Ka-band. However, the spacecraftspeed is under 100 km/s for allbut 2 days of the orbit, and thecorresponding angular separa-tion is approximately 1.3beamwidths at Ka-band andone-half beamwidth at X-band.



subsystem utilizes the LGAs and the MGA toaccommodate communication with the Earthduring cruise. The LGA and MGA designs haveprior flight heritage.

Two LGAs were incorporated into the designfor emergencies that could occur early in the mis-sion or when mission events preclude pointingof the MGA and HGA. The LGAs are mounted180° apart; each provides angular coverage of45° (±22.5°) and peak gain of 10.75 dBi at theX-band downlink frequency, and, for a 5-Wtransmitter, a telemetry rate that exceeds 10 bpsinto a 70-m ground station for spacecraft-to-Earth distances of less than 2 AU. Each allows acommand rate of at least 31.25 bps with over a6-dB margin under the same conditions on theuplink.

The primary antenna used during the cruise por-tion of the mission is the MGA, which is a singleMGA with angular coverage of 10° (±5°) andgain of 20.4 dBi at the X-band downlink fre-quency. For an onboard 5-W transmitter, the linksupports a data rate of 10 bps with 3-dB margininto a 70-m ground station at spacecraft-to-Earthdistances up to 6 AU (the maximum distance

Consequently, an X-band command link was se-lected. The uplink allows a command rate of500 bps with a link margin of 16 dB when space-craft speed is 100 km/s. Alternatively, separate34-m assets could be used for Ka-band telem-etry and commanding, but this option requiresthat the spacecraft carry a Ka-band receiver,which is less desirable because of technologicalimmaturity and lack of heritage in interplanetarymissions. Besides, the low data rate associatedwith commanding diminishes the need for Ka-band. The X-band configuration chosen utilizesreceiver hardware with strong flight heritage fromthe TIMED and CONTOUR programs.

4.4.3.2. Baseline Implementation

The selected telecommunications architecture isshown in Figure 4-21. The telecommunicationsubsystem features bidirectional communica-tions at X-band through all of the spacecraft’santennas: two LGAs, one MGA, and one HGA.In addition, it includes a telemetry link at Ka-band via the HGA. Although the HGA affordsthe best link, it sets an upper bound for the totalpointing error relative to the spacecraft-to-Earthline of sight at 0.2°. The telecommunication

Figure 4-21. Telecommunications architecture.

IEM B

Freq.Mult. ◊4

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02-0817R-24



from Earth in the current mission design). Un-der these circumstances the command rate is atleast 125 bps with over 6-dB of link margin. Thebaseline design facilitates MGA pointing byhaving the MGA mounted on the gimbal thatpoints the HGA.

The HGA is the preferred aperture for commu-nications transactions during the solar encoun-ter and data retrieval phases of the mission,including all of the science telemetry. The mainreflector is 0.8 m in diameter and utilizes a di-chroic subreflector to transmit a right-hand cir-cularly polarized wave at Ka-band. A hornbehind the subreflector provides bidirectionalcommunications at X-band. Umbra size (and ul-timately lift mass), constrains the size of themain reflector and the focal length of the optics(f/D ≤ 0.4).

Figure 4-21 shows how the antennas just de-scribed connect to the spacecraft electronics. Thespacecraft’s transmitter section consists of a pairof downlink cards, each within its own integratedelectronics module (IEM); four solid-state poweramplifier (SSPA) modules, two at X-band andtwo at Ka-band; a pair of ultrastable oscillators(USOs); and a distribution network to couplesignals between the SSPAs and any of the avail-able antenna apertures. We envision that primepower is applied to only one of the four SSPAsto enable either X-band or Ka-band communi-cations on the telemetry link. Both frequencybands are serviced by block-redundant SSPAs.The output from either of the X-band SSPAs maybe steered to any of the antennas through a net-work of single-pole-double-throw (SPDT) andtransfer (XFER) switches, which are themselvesconfigured for redundant operation. A frequencyquadrupling circuit drives the input of each bankof Ka-band SSPAs.

The downlink card starts with the same designbasis as those flown on TIMED and CONTOUR.However, it must be enhanced to support turbo-encoding to take advantage of the reduced

signal-to-noise ratio (S/N) at which these DSMSlinks can operate.

The uplink card is the same as that under devel-opment for the New Horizons mission. Thetransceiver was selected based on the uplink anddownlink cards rather than a transponder be-cause of the ease of integration brought aboutby consolidating the baseband and RF modula-tion functions within the same enclosure as theIEM. The USOs are also based on CONTOUR/New Horizons heritage; we start with that de-sign and provide enhancements that lower massand prime power consumption and improve im-munity to irradiation.

4.4.3.3. Noncoherent Navigation

Instead of standard Doppler tracking tech-niques using a transponder, Solar Probe’scommunications system includes features formaking accurate Doppler measurements witha transceiver. This approach eliminates the needto carry a transponder. For an RTG-poweredspacecraft such as Solar Probe, this design pro-vides a significant and needed reduction in re-quired spacecraft power.

An approach was developed and demonstratedon CONTOUR that allows accurate Doppler mea-surements to be made with a simple transceiverand a moderately stable oscillator (Jensen andBokulic 1999). The approach involves a measure-ment of the ground station’s uplink carrier fre-quency by the spacecraft, which provides resultsto the ground station via telemetry. The groundstation combines the information from telemetrywith measurements of the spacecraft’s downlinkcarrier frequency to obtain a Doppler measure-ment that has the same precision as one obtainedwith a coherent transponder. A conceptual draw-ing of the approach is shown in Figure 4-22. Thisnavigation approach to Doppler tracking with atransceiver does not require any new equipmentat the ground stations. The technique impactsground operations because the observed Doppler



frequencies must be corrected by telemetry priorto their use in orbit determination. Extensive tests(Jensen and Bokulic 2000) demonstrated that thisnoncoherent telemetry-assisted technique makesvelocity measurements with a precision of≤0.1 mm/s for 1-min measurement intervals. Thususe of a lighter-weight, lower-power-consump-tion transceiver yields results equivalent to those

obtainable from a coherent technique with atransponder.

4.4.3.5 Summary

The telecommunications subsystem meets all themission requirements. The continuous highdownlink data rate throughout the perihelionpassages is critical to ensure a successfulscience mission and is accommodated byimplementing a Ka-band downlink. A uniqueadvantage is the opportunity for telemetry ofCruise-Mode science data return during the con-tacts for noncoherent navigation. Some technol-ogy development is needed to implement ahigh-efficiency Ka-band system; the work planis detailed in Section 5, Risk Mitigation.

4.4.4 Power


Requirements that drive the design of the powersubsystem are set by the extremes of the SolarProbe orbit and its dual solar encounters (Table4-1, requirements 1, 3, 14-16). The impacts onthe power subsystem design are summarized inTable 4-11.

4.4.4.2 Primary Power Source Selection

Several candidate sources for primary powerwere considered. Solar power was ruled out earlyin the trade-offs because of the mission’s wideextremes in the solar environments. Currently,no spacecraft with solar panels has flown be-yond the 2.1 AU reached by NEAR, and nospacecraft has been conceived to fly closer than

Table 4-11. Power subsystem requirements summary.

Requirement ImpactProvide reliable spacecraft power to the spacecraft over the entire missionExtremely low solar flux at 5.0 AU aphelionIntense solar flux at 4-RS perihelion

Solar panels impractical; chose multi-mission radioisotope thermal genera-tor (MMRTG) as primary power source

Multiple orbits and solar encounters encompassing over 7 years design lifeMaximum average power required, driven by the solar encounter of 248 W

Three MMRTGs for end-of-missionpower level

Maximum peak power required, driven by the solar encounter when propul-sion is active (momentum management) or when the reaction wheels areproviding large torques to slew or compensate for dust impacts. The totalpower could theoretically peak to 458 W

Lithium ion battery secondary powersource for extended peak loads

Figure 4-22. Two-way, noncoherent navigationconcept.

02-0817R-25



the 0.3 AU planned for MESSENGER. SolarProbe will spend about 10 days inside of 0.3 AUduring each solar encounter. On the basis of thepower required and the mass constraints, wedetermined that using batteries alone for powerduring this period would be impractical.

The only other mature space power sources areradioisotope sources. The RTG design that hasbeen flown for the past 25 years is no longerproduced and will not be available to supportSolar Probe. However, plans are for two radio-isotope sources to be available to support a 2010Solar Probe mission: the multi-mission RTG(MMRTG) and the Stirling radioisotope genera-tor (SRG).

The MMRTG is has essentially the same basicdesign as previously flown RTGs but is smaller,carrying eight general purpose heat sources(GPHSs) rather than 18. Transfer of the heatgenerated by the radioisotope units to electricalenergy is essentially the same as in previous RTGdesigns. Each MMRTG is specified to provideat least 110 W BOL (beginning of life) space-craft power per unit.

The other possibility, use of SRG power sources,offers the advantage of significantly better effi-ciencies than RTGs and would require less plu-tonium to supply the same power. SRGs arecurrently being developed at NASA Glenn Re-search Center, and development is scheduled tosupport a 2010 launch. Currently SRG designshave no flight heritage. However, developmentand testing of prototype units at NASA Glennlook promising. An SRG is baselined for theMars 2007 Lander.

The MMRTG option was selected as the cur-rent baseline, primarily because the technicalrisk is lower than that for the SRG option. TheMMRTG design is basically a repackaged ver-sion of flight-proven RTG technology. In addi-tion, we had some concerns about whether theSRG, with its moving electronic parts, would

meet the electromagnetic compatibility and in-terference (EMC/EMI) requirements of the sen-sitive plasma wave instrumentation that SolarProbe will carry. This decision can be revisitedin the future and reevaluated as the MMRTGand SRG designs mature.

4.4.4.3 Secondary Power Source Selection

Because potentially large transient peak powerloads were identified during the study, a second-ary power source was considered necessary.High transient peaks will come primarily fromreaction wheels, thrusters, and some of the in-struments. Most of these peaks are less then 1 sin duration, but during momentum managementand �V maneuvers, significant power drawscould last for several minutes.

These potential peak power loads are likely tolast too long for reliance on a capacitor bankalone, which is typical for RTG-poweredmissions. A small battery, on the order of 4.5A-h, would be more than adequate to handlethese transient peak power loads. A lithium ionbattery was chosen based on its high energy den-sity and predicted availability in the next fewyears. Lithium ion power sources have beenunder significant development in recent years,and the risk of going to this technology was con-sidered low.

4.4.4.4 Power System ImplementationDescription

A block diagram of the power system compo-nents is shown in Figure 4-23. The system con-sists of three MMRTGs, a shunt regulator unit(SRU), external shunts, battery charger, boostconverter, battery, and a power distribution unit(PDU). The three MMRTGs provide totalpower of 330 W BOL. Power available decaysto 274 W at the end of the mission. Figure4-24 shows power decay from earlier RTGmissions.



Figure 4-23. Power system block diagram showing one of the three MMRTGs.

ExtrenalShuntRresistor

+

Bus Caps

Shunt Driver

-

Regulated Power Bus

Load+RTG

Shunt

Bus Controller

-

LoadSwitchController

CommandDecoder

1553 -Interface

I - Limit

SRU PDU

BoostConverter

BatteryCharger

Li IoBATTERY

02-0817R-26

RTG power will also degrade some duringlaunch. On the launch pad the unit is cooled withforced air. At the instant of liftoff, the airflow stops and the unit is enclosed in the launchvehicle fairing for much of the lift phase. As a

consequence, the radiator plates of the MMRTGsheat up and the power output decreases. Soonafter the fairing is ejected, xenon gas will bevented and the MMRTGs will radiate thermalenergy to space, which will allow the RTG units

1Ulysses power not measured directly, isestimated from equipment power draw.

Voyager 1 RTG 3Mission: 5 Yrs

(Highest)

GalileoMission: 8 Yrs

LES 8 RTG 2Mission: 5 Yrs

(Lowest)

Ulysses 1

Mission: 4.7 Yrs

CassiniMission: 10.8 Yrs

0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26

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er R

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MM

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-196

a

Time, Years

Voyager 2Average of

3 RTGs

Voyager 1Average of

3 RTGsVoyager 1Analytical

Model Prediction

1Ulysses power not measured directly, isestimated from equipment power draw.

Voyager 1 RTG 3Mission: 5 Yrs

(Highest)

GalileoMission: 8 Yrs

LES 8 RTG 2Mission: 5 Yrs

(Lowest)

Ulysses 1

Mission: 4.7 Yrs

CassiniMission: 10.8 Yrs

0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26

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er R

atio

, P/P

o

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RTG

-196

a

Time, Years

Voyager 2Average of

3 RTGs

Voyager 1Average of

3 RTGsVoyager 1Analytical

Model Prediction

02-0817R-27

Figure 4-24. Empirical RTG decay rate data.



to reach normal operating efficiency in approxi-mately 8 hours. The spacecraft design will limitrequired power loads during launch, and initialmaneuvers will be conducted after adequatepower margin is available.

The SRU controls power to the spacecraft usingresistive shunts to dissipate unneeded power anda small capacitor bank to provide an even flowof current to the spacecraft. As discussed earlier,a lithium ion battery provides secondary powerto accommodate peak loads. The battery ischarged and controlled by the battery chargerwhen there is surplus power, so that the batteryis kept fully charged throughout the mission.When the load power exceeds the MMRTGoutput, the bus voltage drops. The boost convertermonitors the bus voltage, and at 29 V it begins totake energy from the battery to keep the bus volt-age at 29 V. The PDU provides relays andswitches controlling loads to various subsystems.

4.4.4.5 Experience with RTG-PoweredSpacecraft

An RTG-powered spacecraft must cross signifi-cant hurdles to get approval for flight. The

National Environmental Policy Act (NEPA)process to approve RTG-powered spacecraftlaunches is very involved. Nonetheless, RTGsare used for a number of space missions. JHU/APL has experience in obtaining launch ap-proval on multiple projects that used RTGs andis quite familiar with the NEPA process. Weare currently developing the required documen-tation for the New Horizons project, and wehave assisted/developed the NEPA documentsfor Aerospace Nuclear Safety Programs since1971. These include Voyager, Pioneer, Galileo,Ulysses, Mars Pathfinder, Cassini, Mars 2001and Mars 2003. JHU/APL’s previous experi-ence with the NEPA approval process is sum-marized in Table 4-12.

4.4.4.6 Summary

The power subsystem meets all requirementsderived for this mission with a low technical riskfor implementation. The availability of theMMRTG is governed by activities beyond thescope of the Solar Probe project, as detailed inSection 5, Risk Mitigation. However, the cur-rent development schedule the Department ofEnergy has released for the MMRTG meets our

Table 4-12. Previous JHU/APL contributions to NEPA launch approval process.

MissionFlightRTGUnits

No. Radio-isotope

Heater Units

YearLaunched JHU/APL Contributions

Galileo F-1,F-4

Over 100 1989 Response analysis of light-weight radioisotope heater units(LWRHUs) to inadvertent Earth reentry during Venus-Earth-Earth gravity assist (VEEGA); technical review of GeneralElectric (GE) analysis; arc-jet tests of GPHS, graphite impactshell (GIS), and LWRHU to simulate reentry.

Ulysses F-3 1990 Technical review of GE analysis; wind tunnel tests of GPHS,GIS, and LWRHU.

MarsPathfinder

N/A 3 1996 Response analysis of LWRHUs to solid rocket motor (SRM)fires.

Cassini F-2, F-6, F-7

117 1997 Response analysis of LWRHUs to inadvertent Earth reentryduring Venus-Venus-Earth-Jupiter gravity assist (VVEJGA);preliminary GPHS reentry safety assessment, GPHS motionstudies; technical review of Lockheed Martin Astronauticsanalysis; answer Interagency Nuclear Safety Review Panel(INSRP) inquiries, launch contingency effort.

Mars 2001 N/A 3 2001(landercancelled)

Issued nuclear risk assessment for draft environmentalimpact statement (DEIS); issued preliminary safety analysisreport (PSAR) (Tetra Tech NUS, Inc. [TtNUS]); tested SRMfires.

Mars 2003 N/A Up to 22 total 2003(planned)

Issued “quick-look” nuclear risk study; reviewed databook(TtNUS); reviewed STAR 48 breakup system design.



2010 launch date schedule. The lithium ion bat-tery technology is maturing rapidly and is cur-rently being used in other space applications. Itis expected to be one of the standard battery tech-nologies available within the Solar Probe devel-opment time frame.

4.4.5 Avionics Architecture

The Solar Probe avionics architecture is illus-trated in the system block diagram in Figure 4-21 and described in the following sections.


The requirements that most affect the SolarProbe avionics subsystem architecture are sum-marized in Table 4-13. They result from the needto ensure successful science mission against therisks and uncertainties in the extreme spacecraftenvironment (Table 4-1, requirements 1, 3, 23,and 24).

4.4.5.2 Data Bus

The initial selection for the baseline Solar Probedata bus is a 1553B bus. This data bus is astandard configuration with significant flightheritage. Performance of the 1553 bus is morethan adequate to handle spacecraft data traffic,as well as the instrument peak data rates definedin the AO of 112 kbps. As the Solar Probe projectdevelops, this selection will be revisited in a for-mal trade study.

4.4.5.3 Integrated Electronics Module (IEM)

Solar Probe uses an IEM architecture for hous-ing most of the avionics hardware. This approach

is consistent with previous APL spacecraft pro-grams and will take advantage of extensive useof heritage hardware. Each IEM includes theflight processor, SSR, command and telemetrycard, telecommunications uplink and downlinkcards described in Section 4.4.3, and power con-ditioning. Solar Probe accommodates two IEMsfully cross-strapped for redundancy.

Command and Data Handling (C&DH), G&C,and spacecraft fault protection functions will beperformed in a single (240-MIPS) radiation-hardened RAD750 processor. The system archi-tectures developed for the CONTOUR andMESSENGER spacecraft provides significantheritage for developing the Solar Probe IEMelectronics.

Each Solar Probe IEM will include a 128-GbitSSR board based on the 64-Gbit data recordercurrently being developed for the New Horizonsmission. The New Horizons SSR design uses 1-Gbit flash memory integrated circuits that arephysically stacked to form each of the 16 inde-pendently powered 4-Gbit memory banks. TheNew Horizons SSR board control logic enhancessystem fault tolerance and survivability by main-taining an EEPROM map of bad memory blocksand substituting functional memory blocks forfailed portions of the SSR memory. JHU/APLsuccessfully tested the 1-Gbit flash memoriesat the New Horizons SSR total dose limit of40 krad while operating the memories at a 10%duty cycle.

The proposed Solar Probe SSR board will formsixteen 8-Gbit memory banks by stacking four(currently available) 2-Gbit flash memories.Minor modifications to the New Horizons SSRflash memory control field-programmable-gate-array (FPGA) logic will be required to accom-modate the larger memories, but the SSRarchitecture and most of the control electronicswill be essentially unaltered. Tests will be con-ducted to verify proper operation of the 2-Gbitmemories at a total dose limit of 100 krad while

Table 4-13. Avionics subsystem architecture re-quirements summary.

Requirement ImpactMaximize science returnduring each solar encounter

Redundant 128-GbitSSRs

Provide a fault-tolerant ar-chitecture with adequateredundancy to ensure mis-sion success

Dual IEM plus dualAIU with backup at-titude control function

Survive high-radiation envi-ronment introduced by theJupiter flyby

Qualify to 100 kradtotal dose



operating at a 10% duty cycle. Since the SSRboard does not have to be powered during theJupiter flyby, turning the SSR off near Jupiterremains an option if additional (total dose) mar-gin is required. An alternative Solar Probe SSRdesign that stacks eight 1-Gbit memories to formeach memory bank can be considered if totaldose testing of the 2-Gbit flash memories is nottotally successful.

During a solar encounter, both IEMs will bepowered and each IEM will process uplink com-mand messages, detect and correct spacecraftfault conditions, and record all instrumentscience data (i.e., redundant recording of all sci-ence data). The primary (selected) IEM proces-sor will perform all G&C control functions andwill serve as the spacecraft data bus controller. Ifa critical fault condition is detected in the primaryIEM, data bus controller and G&C functions canbe quickly switched to the redundant IEM pro-cessor without rebooting. During non-critical mis-sion periods, the redundant IEM may be turnedoff to reduce spacecraft power bus loading.

4.4.5.4 Attitude Interface Unit (AIU)

Redundant AIUs housed in a single chassis pro-vide all required interfaces between the 1553Bdata bus and the attitude system sensors (horizonsensors, digital solar aspect detectors [DSADs])and actuators (reaction wheels, thrusters). Unlikethe IEM processors, the AIU processors will op-erate in a standby redundant mode, with only oneunit powered at any time. The selected AIU servesas a backup attitude control processor, running asimplified control algorithm that will take overattitude control functions should a critical atti-tude error occur during a solar encounter. TheSolar Probe AIUs are similar in concept and com-plexity to the RTX-2010–based NEAR AIUs;each consists of a processor board, dedicated at-titude system interfaces, and power conditioningelectronics. The current AIU baseline design in-cludes a RAD750 processor that offers the ben-efit of a common design with the IEM processor,

but alternative, lower-power processors will beconsidered.

4.4.5.5 Instrument Data Processing

Because no instrument teams were working withus during this study, we had to assume an in-strument data processing architecture. Based onrecent experience on several other JHU/APLspacecraft programs, we assumed that the in situand the remote sensing instrument suites wouldeach have their own DPU acting as a remote ter-minal on the spacecraft data bus. All instrumentscience and housekeeping data and all instru-ment commands will be transferred between theDPUs and the IEMs over the redundant 1553Bspacecraft data bus, which can easily accommo-date the maximum planned instrument data rateof 112 kbps.

4.4.5.6 Summary

The avionics subsystem meets all the require-ments with a robust and low-risk implementa-tion. The issue of radiation hardness of the SSRsis detailed in Section 5, Risk Mitigation. Asidefrom that issue, the subsystem can be imple-mented within current technical capability.

4.4.6 Data Handling Approach


The Solar Probe data handling subsystem is thecoordinated operation of the onboard data storeand telecommunications. The requirements fordata handling are to maximize both the quantityand the reliability of the science data return(Table 4-1, requirements 23 and 24). The hard-ware elements of the data handling system weredescribed in Sections 4.4.3 and 4.4.5. Here wefocus on the operational approach.

4.4.6.2 Data Management Approach

The primary objective of the Solar Probe datamanagement approach was to maximize both thequantity and the reliability of the science data



return. The spacecraft design provides a large,redundant data storage capability as well as alimited-bandwidth, continuous real-time link.

Each of the two 128-Gbit SSRs is sized to holdan entire 20-day solar encounter’s worth ofplanned science data with margin. Although So-lar Probe maintains a continuous real-timedownlink during the encounter, this link plays abackup role, ensuring a scientifically significantdata return even in the event of a catastrophicfailure in the harsh near-Sun environment. Theredundant recorders capture all science data, andreal-time data paths and redundant X- and Ka-band downlink capability ensure maximumprobability of returning prioritized mission sci-ence data.

The encounter science data collection profile isbroken into three phases as follows:

• Primary: –10 days to –1 day; +1 day to+10 days

• Critical II: –24 hours to –8 hours; +8 hoursto +24 hours

• Critical I: –8 hours to +8 hours

The total data rate increases as Solar Probe ap-proaches the Sun and decreases as the space-craft recedes, reflecting the criticality anduniqueness of the in situ coronal measurements.The distribution of the data rates between theinstrument suites and within the suites is opti-mized to the measurement requirements of eachphase. Since the entire 20 days of encounterdata are stored in each SSR, the data storagecapacity is driven by the instrument data raterequirements. For this study, the requirementswere developed from a JPL-provided summaryof the instrument AO responses. The stored datarate profile is shown in Figure 4-25. The SSRcapacity of 128 Gbits offers ample space forscience and spacecraft housekeeping data takenduring the encounter.

The real-time data rate is determined by the in-stantaneous capability of the downlink. This rate

is affected by many factors including the Sun-Earth-Probe geometry, Earth–Probe distance, el-evation angle at the ground station, weather at theground station, and the downlink band. Hence, themission operations staff can adjust the real-timedata rate to maximize the return depending uponthe current operating conditions. The spacecrafthas the capability to maintain several science datarate tables corresponding to different real-time datarates. The nominal real-time science data collec-tion rate profile is shown in Figure 4-26. Any un-used real-time bandwidth is automaticallyallocated to playback of the data recorders.

Starting 3 days after perihelion, 2 hours of thereal-time link each day are devoted to playbackof the most critical encounter data. At the endof the encounter, a full recorder playback is ini-tiated using 16 hours per day of DSMS contacttime. Figure 4-27 shows the overall data profilefor each encounter. The stored data total reflectsthe required instrument data rates depicted inFigure 4-25. The total science return line reflectsthe sum of the real-time data transmitted duringthe encounter plus the stored science data placedin the recorder. Additional science data can beaccommodated, limited only by total capacityof the recorders and the extended operation costfor downlink time.

The onboard data handling approach places theonus for instrument data processing on the in-strument DPUs. Once generated, the instrumentdata are sent to the DPU for that instrument suite.The DPU performs any required data evaluationand compression. Low-rate science data areplaced into CCSDS frames and sent to theC&DH component over the 1553 bus. Aftercompression, high-rate image data are convertedinto CCSDS frames and buffered in the DPU tobe sent over the 1553 bus as well.

4.4.6.3 Summary

The data handling subsystem meets require-ments and exceeds the expected data return over



Solar Probe Perihelion Passage Science Data Storage

0

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rag

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ata

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All-sky 3-D Coronagraph Imager

Magnetometer

Solar Wind Ion Comp. & Electron Spec.

Energetic Particle Composition Spec.

Plasma Wave Sensor

Fast Solar Wind Ion Detector

02-0817R-45

Figure 4-25. Solar encounter instrument stored data rates.

Solar Probe Perihelion Passage Real-Time Science Data

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Energetic Particle Composition Spec.

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Fast Solar Wind Ion Detector

02-0817R-29

Figure 4-26. Nominal real-time science data collection profile.



the 1999 mission description by a factor of 3.This can be accomplished using establishedhardware described in Sections 4.4.3 and 4.4.5with little overall technical risk.

4.4.7 Guidance and Control


Pointing Requirements. The pointing require-ments for Solar Probe come from two basicsources. The first source, the AO documentation,defines the following pointing requirements tosupport the remote sensing instruments duringsolar encounter (Table 4-1, requirements 19–22):

• Absolute pointing requirement ≤0.3°• Absolute attitude knowledge ≤0.05°• Jitter ≤0.005° in 1 s

For this study we assumed that the nominalboresight was solar nadir and that these require-ments would be applied over the entire solar en-counter except during brief periods for necessary

momentum management. Refinement of theserequirements and assumptions will require in-teraction with the science team.

The second source of pointing requirements isallocated based on the pointing requirements ofthe HGA. Selection of Ka-band transmission andthe limit of a 0.8-m HGA requires a pointingaccuracy of 0.2°. Accounting for misalignmentsand ability to calibrate the antenna on orbit, theattitude control system was allocated 0.025° in-ertial knowledge accuracy and 0.05° for attitudecontrol.

Environmental Requirements. Several environ-mental factors were drivers for the Solar ProbeG&C design. First, as the spacecraft approachesperihelion, a star tracker looking away from theSun through the solar corona will see sunlightreflected off dust particles (Klaasen 2002). Inaddition, the coronal brightness may vary by afactor of 10 due to “local condensations,” and

Solar Probe Science Data Return

0

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-10 0 10 20 30 40 50

Time from Perihelion (days)

Dat

a (G

b)

Realtime Science

Stored Science Data

Total Science Collected

Total Stored Data

Current Stored Data

02-0817R-28

Figure 4-27. Solar encounter data collection and storage profile.



+X

+Y

+Z

02-0817R-30

coronal lighting reduces the S/N ratio on thecharge-coupled device (CCD) in the star tracker,thereby reducing the number of detectable starsand degrading the performance of the startracker. These phenomena considerably limitedthe choice of star trackers and also drove thedesign to operate two star trackers throughoutthe solar encounter.

Next, solar pressure will be very high and willchange rapidly during the solar encounter. Dur-ing this phase, the cone-shaped heat shield is tobe pointed toward solar nadir. Because the cen-ter of photon pressure is ahead of the center ofmass, the solar pressure torque is destabilizingand becomes an important part of the dynamicsof the spacecraft near perihelion. The solar pres-sure torque is 0.08 N⋅m/radian (0.0014 N⋅m/de-gree) at perihelion and decreases with distancer from the Sun as 1/r2. Initial analysis (Lisman2001; Tarditi et al. 2001) shows that without con-stant attitude control at perihelion, the photontorque causes the attitude to change from 0.2°to 3° in 3 min, thereby exposing the instrumentsand the spacecraft bus directly to theintense solar flux. In addition, even relativelysmall misalignments of the heat shield inducessignificant torque and momentum buildup, po-tentially forcing more frequent use of thrustersfor momentum management.

Solar dust impacts were also an important atti-tude control consideration. Data extracted fromMann (2001) and Lisman (2002) indicate a 1%chance of encountering a grain 100 �m or largerfor each square meter of surface area projectedinto the relative velocity direction. The dust par-ticles are mostly in a nearly circular orbit within±30° of the solar equator and are in greater con-centration near the equator. The momentumand attitude can be controlled by wheels for<130-�m dust and by warm and ready thrustersfor <315-�m dust. There is significant uncer-tainty in the dust model due to the paucity of insitu data, and so a conservative design is waspursued.

Finally, the large velocity of the spacecraft re-sults in significant stellar aberrations due to rela-tivistic considerations. Relativistic effects willcause apparent shifts in the Earth position, solarnadir, and stellar references, requiring correc-tion terms as part of the G&C design.

4.4.7.2 Functional Description

The Solar Probe spacecraft coordinate frame isillustrated in Figure 4-28 and is defined asfollows:

• Origin centered at base of adapter ring• Z-axis through center of adapter ring and apex

of heat shield• –Y-axis parallel with nominal boresight of

antenna and in the plane of the bottom of theadapter ring

• X-axis formed using right hand rule

Shortly after deployment, the G&C system willneed to despin the spacecraft from 60 to 0 rpm.This spin rate is induced by the spin-stabilizedStar 48B stage. Despin will be performed byonboard thrusters until the total angular mo-mentum is reduced to 0.1 N�ms. Then thewheels take over and slew the spacecraft to thedesired attitude.

For most of the mission the spacecraft will main-tain 3-axis pointing control with the –Y axis

Figure 4-28. Solar Probe coordinate frame.



(nominal HGA/MGA boresight) pointed towardEarth. This attitude can be maintained when thespacecraft at least 0.8 AU from the Sun. Whilein this cruise attitude mode, a slowrotation of 0.5 rpm will be introduced to reduceangular momentum buildup from solar pressuretorque. Occasionally, momentum will build upenough that momentum dumps using thrusterswill be needed. Adequate pointing control willeasily be maintained to keep the MGA pointedat Earth to within 5° circular error 3�.

As the spacecraft comes inside of 0.8 AU, it mustchange attitude so that the TPS points towardthe Sun and keeps the instruments and sub-systems within its protective umbra. Some off-pointing from solar nadir may be necessary tokeep the HGA and MGA pointed at the Earth,as long as sensitive instruments and subsystemsare not exposed to the Sun.

During the solar encounter (P ± 10 days), thespacecraft attitude will be maintained so that theZ-axis is pointed toward solar nadir and the HGAand science pointing requirements are main-tained. As a result of the intense solar pressures,momentum dumping will be much more fre-quent during this period. For short periods,thrusters will fire to remove angular momentumand the requirements for instrument pointing andcontrol may not be maintained. Each momen-tum management maneuver will be completedin under 1 min.

An alternative approach is being considered forthe periods of most intense solar pressure; thisapproach involves dynamically controlling thespacecraft so that the photon torque can actu-ally be used to control momentum buildup. Suchan approach requires an intentional offset in theheat shield that is controlled automatically bythe feedback control system. If this approach isimplemented, traditional momentum dumpingusing thrusters will not be needed during thisperiod. Implementation will require additionalanalysis and interaction with the science teams

to ensure that the intent of the pointing require-ments can be met.

4.4.7.3 Hardware

Attitude Determination. Attitude determinationwill be performed by redundant star trackers anda high-precision IMU. The baseline star trackerselected was the Sodern SED-16. This startracker meets performance, radiation, and de-sign life requirements, and preliminary data in-dicate that it will perform adequately in thepresence of coronal lighting near the Sun. Thestar tracker pointing directions (boresights in theX-Y plane and Y-Z plane at 45° from the –Y axis)were chosen to ensure that the star trackers viewdifferent parts of the solar corona at all times,minimizing the chances of both trackers beingblinded simultaneously by locally high coronalbrightness.

The IMU initially selected as the baseline is theNorthrop Grumman Space Inertial ReferenceUnit (SIRU). This IMU is a derivative of previ-ous IMUs that have significant flight heritageand is designed to meet high radiation and de-sign-life requirements associated with deep-space missions. Although the SIRU gyro hasexcellent drift stability, the drift itself can be upto 2° per hour estimated to an accuracy of 0.05°per hour. In the absence of star tracker updatesand a reliable bias estimate, safe attitude con-trol can be maintained for up to only 20 min.The SIRU unit has 4-for-3 redundancy for bothgyros and accelerometers and redundant elec-tronics and interfaces.

To accommodate a safing attitude determinationsystem during cruise, digital Sun sensors aremounted to provide almost 4� steradian cover-age and allow an attitude reference relative tothe Sun if an attitude determination anomalyoccurs. At distances beyond 3.0 AU, simplypointing the –Y-axis toward the Sun allows con-tact with the Earth within the MGA beamwidthof 10°. Inside of this solar distance, a small step



star search pattern will be performed to reestab-lish contact with the ground.

Because of possible long-duration star trackerblinding, system resets, or other attitude controlanomalies, we thought it prudent to baseline anew solar horizon sensor for safing during thetime that the spacecraft needs to be protectedbehind the TPS umbra. One such concept beingevaluated is a device mounted to the end of theboom. The detector consists of a conical ring ofcarbon–carbon material, a mirrored conical re-flector, and a detector array with pinhole lens.The detector array resides in a small electronicsbox, which contains readout electronics for boththe detector and a set of thermistors.

If an attitude error reaches a designated thresh-old, the edge of the conical ring is illuminatedand projected on to the detector. The processedsignal is used to provide attitude control for safingpurposes during the solar encounter. The detec-tor could potentially be based on a micro-DSADcurrently being developed for the New Horizonsmission. Another option being considered placessensors along the outer ring of the TPS.

Attitude Control. Primary attitude control willbe performed using four reaction wheelsmounted to provide 4-for-3 redundancy. Due topower constraints, only three wheels will bepowered and the fourth carried as a cold spareduring the solar encouter. Ithaco TW-4B200wheels were initially selected to meet the mo-mentum storage and torque requirements to off-set dust impacts up to 1.5° off solar nadir withinthe tight power and mass constraints imposedby the mission.

An alternative approach, use of all-thrusterdead-band control, was also considered. Thisapproach uses small minimum-impulse-bitthrusters like those used on Voyager and

Cassini, eliminating the need to carry reactionwheels and significantly reducing the spacecraftpower requirements. The power reductionscould be significant enough to eliminate thebattery and charging system, since transientpower peaks would be reduced considerably.Preliminary analysis indicates that pointing re-quirements might be met with thruster control.The reduction in reaction wheel mass is mostlyoffset by the increase in mission propellant.This approach was not baselined because thecurrent assumption is that 2.5-mg/s outgassingrequirement encompasses everything on thespacecraft including thruster exhaust. A singleimpulse bit using two coupled thrusters expends11 mg over a 400-ms duration. The thruster con-trol option will continue to be explored onceadditional inputs from the science communityare available to address the outgassing issue inmore detail.

Sixteen 4-N thrusters will still be carried for �Vmaneuvers and momentum management. Thethrusters have been initially placed to provideredundant coupled pairs about the spacecraftcenter of mass, providing the capability to ma-neuver efficiently in any direction while main-taining contact with the ground.

4.4.7.4 Summary

The G&C subsystem meets requirements witha low-risk functional approach and mostly ex-isting hardware. A conservative design over-comes the significant uncertainties in theproperties of the unexplored environment thatSolar Probe encounters in the solar corona. Bet-ter environmental knowledge would enable re-finements in the concept and a better estimateof the required resources. Section 5, Risk Miti-gation, details a prudent plan for reducing envi-ronmental uncertainties and developing a newsolar horizon safing sensor to add robustness tothe design.



4.4.8 Propulsion


Propulsion requirements are derived from themission design and G&C requirements (Table 4-1, requirements 1–3 and 19–22). The function ofthe propulsion system is to provide thrust forperforming �V maneuvers and managing angu-lar momentum. A thrust level of 4 N was consid-ered adequate to balance the needs foracceleration during maneuvers, minimum-impulse-bit control to maintain attitude duringmomentum dumps, and significant margin intorque to overcome expected dust particle impactsduring the solar encounter. The initial �V esti-mate of 225 m/s and attitude control budget of6 kg requires a total available propellant load of69.1 kg. Sixteen thrusters were needed to pro-vide adequate redundancy and flexibility tomaintain contact through the MGA for any �Vmaneuver.

4.4.8.2 Alternatives Considered

Several propulsion system types were considered,including both electrical and chemicaloptions. All electric propulsion systems were im-mediately discarded, however, because of the lim-ited power available on the Solar Probespacecraft. A simple blow-down monopropellantsystem that is fully compliant with the SolarProbe requirements was ultimately selected,though a dual-mode system with bipropellant �Vthrusters and monopropellant attitude controlthrusters might theoretically weigh less. This se-lection was made for a variety of reasons, includ-ing cost, reliability, and packaging advantages.(For example, the bipropellant engines are largeand radiatively cooled, and the pointing require-ments of the spacecraft combined with the um-bra constraints essentially preclude their use.)

4.4.8.3 Concept Description

The baseline propulsion system architecture isshown in Figure 4-29. This design is similar in

architecture to almost every hydrazine propul-sion system flying today. Sixteen thrusters pro-vide forces in all required directions, and eachthruster has series-redundant control valves toprotect against leakage. These thrusters aregrouped into two redundant sets, which provide�V in groups of four and can be used singly orin groups of two to four for momentum man-agement. Hydrazine propellant and nitrogenpressurant are stored in a single tank whose pres-sure decreases as propellant is depleted.Pressurant is separated from the propellant byan elastomeric diaphragm within the tank. Latch-ing valves isolate the thrusters from the tank forground safety and system reliability (i.e., in caseof a thruster leak), while manual service valvesare used for testing and loading the system onthe ground. The system’s surge suppression ori-fices keep transient pressures within appropri-ate levels, and the pressure transducers are usedtogether with temperature telemetry to gaugepropellant and monitor system performance inflight. Spacecraft ambient temperatures will bemaintained such that the propulsion system re-quires no heaters except those on the thrustercatalyst beds.

Several flight-proven options exist for each com-ponent of the Solar Probe propulsion system,resulting in a high level of confidence that noqualification testing will be required for theproject. A representative set of heritage compo-nents was selected for preliminary performanceevaluations and shows that system requirementscan be met. The representative thruster selectedis the Aerojet MR-111c 4.0-N rocket engineassembly (Figure 4-30). This thruster has an ex-tensive flight and test history, and it meets orexceeds all Solar Probe requirements. Usingperformance numbers typical of this thruster, atotal maximum required usable propellant loadof 69.1 kg can be calculated for the mission. Thisnumber, together with the range of inlet pres-sures for which the thruster has been qualified,results in a required a tank volume of at least 90



Figure 4-30. Aerojet MR-111c 4.0-N rocket engineassembly, shown with axial nozzle.

Figure 4-29. Solar Probe propulsion architecture.

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02-0817R-31

liters; at least two commercially available flight-proven tanks exist in this size range, affordingflexibility when the final flight propulsion sys-tem configuration is selected.

4.4.8.4 Summary

The propulsion subsystem meets its derived re-quirements with entirely off-the-shelf hardwareand represents a very low risk approach.

4.4.9 Flight Software


The software supports meeting mission require-ments as allocated by the mission, power, ACS,C&DH, communications, and fault protectionsubsystem areas.

4.4.9.2 Software Development Approach

The Solar Probe spacecraft flight software willhave the benefit of significant heritage fromMESSENGER and numerous other missions, in-cluding JHU/APL’s NEAR, TIMED, and CON-TOUR. As on MESSENGER, each IEMcontains a single processor to carry out theC&DH, G&C, interface communications, andautonomy functions. As for TIMED and NEAR,a separate AIU (Attitude Interface Unit) providesan added level of fault protection during the so-lar encounter. As was planned for CONTOUR,both processors will be active during the solarencounter to ensure the collection of all datafrom the two instrument suites.

The five major functional areas addressed byflight software are described below:



• C&DH functions include support of CCSDSprotocols for uplink and downlink, commandprocessing (including macros and time-delayed commands), telemetry processing,Mil-Std-1553 bus management, autonomyrules (fault detection and recovery [FDR],management of redundancy, and otheroperations), SSR management, science datahandling (scheduling, sequencing, and com-pression), thermal management, memoryscrubbing, and other FDR algorithms.

• G&C functions include attitude estimationand control via reaction wheels and thrust-ers, guidance, Mil-Std-1553 bus processing,G&C FDR, memory scrubbing, and telem-etry reporting.

• Interface communications within the space-craft are via the 1553 bus and PCI bus. The1553 bus is used to communicate between theIEMs and major spacecraft subsystems (in-cluding the instrument DPUs, PDU, RF trans-mitters, and the AIU). In addition, the 1553bus is used to cross-strap the two IEMs tooptimize redundancy and increase reliability.The PCI bus is used to communicate with theSSR, command and telemetry card, uplink anddownlink cards, and the instrument interfacecard. Data flow from the instruments to theSSR in both IEMs.

• Autonomy will be critical to carrying out theoperational and safing functions. Sequencesof commands stored onboard can be triggeredat a specified time or when a specific eventoccurs. The use of rules and stored commandsequences eliminates the need for muchspecial-purpose software and decouples theautonomy (operational and safing) algorithmsfrom the software development. Schedulingand sequencing of instruments will beperformed using time-tagged commands andmacros. Instrument state-of-health checkingand other decision-based tasks will be per-formed using autonomy rules.

• Fault protection includes a Cruise Safe Modeand a Solar Encounter Safing Mode. In Cruise

Safe Mode, the spacecraft turns off nonessen-tial hardware and points the antennas towardEarth. In Solar Encounter Safing Mode, theAIU takes control of the G&C functions tokeep the spacecraft pointed toward the Sun.Both modes will be tested independently ofthe operational software.

Solar Probe software will be developed in strictaccordance with the JHU/APL Space Depart-ment software development process. Hardware-in-the-loop simulations using engineeringmodels of the flight IEM are used to test flightsoftware designs and algorithms. These testsaddress each phase of the mission to ensure suc-cess and reduce risk prior to launch. Groundsupport electronics hardware and software aredeveloped to support hardware checkout, soft-ware development, spacecraft integration, andmission operations. Independent acceptance test-ing is performed on mission-critical software.

4.4.9.3 Summary

The flight software will meet all functional re-quirements, and major portions of the designshave significant flight heritage from previousprograms. Software development processes thathave been proven successful on several flightprograms will be implemented on Solar Probe.

4.5 Integration and Test (I&T)

4.5.1 Requirements

Space system I&T has the following objectives:

• Verify system-level performance• Identify unexpected interactions among sub-

system elements• Identify failure modes from design weakness,

material defects, and workmanship• Operate components long enough to identify

failures due to infant mortality• Establish standard operating and contingency

procedures for mission operations



• Verify that the spacecraft will operateproperly through launch and in on-orbitenvironments

4.5.2 I&T Approach

I&T must reduce risk during both system-levelintegration and flight operations. Therefore, amethodical, hierarchical approach, designed touncover potential problems early, must be fol-lowed. This section uses JHU/APL practices andfacilities as an example of such an approach.

Testing starts at the breadboard level where alldesigns undergo interface compatibility testingprior to release for flight fabrication. This prac-tice minimizes the number of interface problemsencountered during system-level integration.Piece parts, components, and boards are envi-ronmentally tested at stress levels higher thanthose the system will encounter in test or opera-tion. Imposing more stressful test levels at lowerintegration levels is a proven technique to findproblems early and minimize problems at sys-tem-level integration. This I&T approach is simi-lar to that used on other JHU/APL spacecraftprograms, including NEAR, TIMED, CON-TOUR, and MESSENGER.

A protoflight approach (qualification levels forflight duration) would be followed for the SolarProbe spacecraft, payload, and third-stage envi-ronmental test programs. Instrument and space-craft components are fully tested, bothfunctionally and environmentally, before deliv-ery for system integration. All components arevibrated in 3 axes at protoflight levels. They alsoundergo operational and survival thermal cy-cling. The design levels used will envelop boththe Atlas V and Delta IV launch vehicle envi-ronments. Results from the structure qualifica-tion testing are then used to correlate thecoupled-loads analysis finite element modelprior to integration.

The Solar Probe spacecraft would be integratedin JHU/APL’s Kershner Space Building. This

spacecraft integration and test facility maintainsclass 100,000, 10,000, and 100 clean rooms. Thebuilding also houses a vibration test facility andthermal vacuum chambers in a variety of sizes.When possible, subsystem and instrument me-chanical models, data simulators, and engineer-ing models will be integrated early withspacecraft structural or electrical components toprovide early mechanical and electrical interfaceverification. Spacecraft integration (Figure 4-31)begins with delivery of the flight-qualified pri-mary structure, which will already include thepropulsion system. The spacecraft harness is theninstalled and rung out to ensure that flight hard-ware can be safely integrated.

A Pre-Integration Review will be held for eachspacecraft component and instrument. The in-tegration team will review the results of the com-ponent or instrument testing program and presentplans for mechanical and electrical integrationwith the spacecraft.

During integration, several special tests will beconducted. These include DSMS RF compatibil-ity tests, time system verification tests, specialguidance and control tests, system self-compat-ibility tests, and mission operations tests.

The three flight RTGs will be integrated afterthe spacecraft is mounted on the launch vehicleat CCAFS. During the spacecraft test program,thermal, electrical, and mass simulators of theRTGs will be used to verify RTG to spacecraftinterfaces and functionality. After integration ofall of the spacecraft subsystems and instruments,the baseline performance test will be performed.Then mechanical alignments will be measuredusing optical cubes mounted on the instrumentsand G&C components.

After the Pre-Environmental Review, the envi-ronmental test program begins. The spacecraftand necessary ground support equipment willbe shipped to NASA/GSFC via air-ride van. Thespacecraft bus will be in one shipping container,



Figure 4-31. Spacecraft integration process.

S/CStructure

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and the TPS will be shipped in a separate con-tainer. The environmental test program startswith the spacecraft vibration tests performed inthe GSFC vibration test facility. The spacecraftwill be in launch configuration (including theTPS) and will be powered during these tests.

Solar Probe will then be moved to the GSFCacoustic test facility for the acoustics and shock/separation tests. These are followed by massproperties measurements and spin-balancing.Finally, the thermal balance and thermalvacuum cycling tests will be conducted. Dur-ing the thermal cycling tests, the DSMS Com-patibility Test Trailer will be utilized to performDSMS RF compatibility tests and end-to-endsimulations of mission operations. Baselineperformance tests will be run prior to the ther-mal vacuum tests and during and after the ther-mal cycling test. In addition, mechanicalalignments will be verified after completion ofthe environmental program.

The spacecraft, TPS, and ground support equip-ment will then be shipped to the KSC launchprocessing facility. Following initial electricaltests, a final baseline performance test will beconducted. The Mission Operations Team(MOT) will be provided time for mission simu-lations and DSMS testing, which will use Mil-71 at KSC.

Next, the flight mechanical build is initiated, in-cluding ordnance installation, flight blanket in-stallation, TPS installation and alignment, andRTG fit checks and installation rehearsals. Af-ter mating with the third stage and installationof the RTG mass models, the final spacecraftspin-balance test will be performed.

Once the spacecraft is moved to the launch padand mounted on the launch vehicle, final prepa-rations are conducted. These include electricalsystem functional testing, launch rehearsals, andflight RTG installation. Readiness reviews and



process rehearsals will be part of a comprehen-sive program to ensure safe and orderly RTGinstallation.

The final steps prior to launch will be the LaunchReadiness Review and red tag item removals.

4.6 Ground and Data SystemsOverview

4.6.1 Requirements

The Solar Probe ground system includes all per-sonnel, hardware, software, data links, and fa-cilities used to conduct I&T operations; toconduct flight tests and operations; to generateand uplink commands; and to receive, process,analyze, and disseminate flight data. The groundand data systems are required to support mis-sion requirements (Table 4-1, requirements 23and 24) as defined by the telecommunications

and data handling implementation approachesas well as the defined concept of operations.Requirements are expected to be better definedonce instrument teams have been selected andcan include their specific requirements. Sincethe study did not have instrument team input,the model was based on previous experience onother recent flight programs.

4.6.2 Design Approach

Figure 4-32 shows the ground systems architec-ture. The ground system architecture is inheritedfrom the MESSENGER and New Horizons mis-sions, which in turn were derived from the low-cost CONTOUR and NEAR systems. Softwarefrom MESSENGER and New Horizons will bereused wherever feasible. The same COTS con-trol center software will be used, as will the JHU/APL-developed telemetry router, server, and

Figure 4-32. Solar Probe ground system architecture.

APL Mission

Operations Center (MOC)

JPL Navigation

APLMission Design and Guidance

& Control

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02-0817R-34



archive system. As on MESSENGER and NewHorizons, SEQGEN software, developed at JPL,will be used for mission planning. Spacecrafthealth assessment and science data processingsoftware will be developed at JHU/APL basedon Solar Probe’s specific requirements.

This ground system architecture is based on acommon, scalable design that supports bench-level testing, spacecraft I&T, and mission op-erations. Using a common set of tools to supportI&T from bench-level testing to integratedspacecraft testing saves both cost and scheduleresources.

The Mission Operations Center (MOC) wouldbe co-located with the MESSENGER and NewHorizons MOCs. Because the MOC groundsystem features the same “look and feel” as thosefor MESSENGER and New Horizons, and be-cause all three spacecraft have relatively com-mon avionics architecture, mission operationspersonnel can be shared. This sharing of per-sonnel reduces cost. Facility equipment such ascomputers, network infrastructure, and periph-erals can also be shared.

A high-fidelity hardware-in-the-loop spacecraftsimulator, maintained in the MOC, will be usedto train controllers and assist in software testing,command load verification, and anomaly resolu-tion. This simulator will be built from spacecraftcomponents that are copies of the flight hardwareand contain the actual flight software. Criticalcommand sequences will be tested on this simu-lator before being uploaded to the spacecraft.

In addition to the hardware-based spacecraftsimulator, a software spacecraft simulator willbe developed to model system resources and testcommand sequences. While the hardware-in-the-loop simulator will operate at real-timespeeds, the software simulator will run manytimes faster, allowing the testing of weeks’ worthof command sequences in several hours.

JHU/APL’s MOC will be connected to theNASA DSMS via leased communications lines.As on MESSENGER and New Horizons, theMOC command workstations, bracketed by net-work firewalls for security, will be connected tothe JHU/APL Space Department network on oneside and the DSMS on the other side. The MOCnetwork will adhere to JHU/APL network se-curity guidelines, the security requirements forconnecting to the DSMS, and any other NASAsecurity procedures mandated for Solar Probe.

Responsibility for many of the data dissemina-tion activities is to be determined, including

• Receiving, archiving, and analyzing all instru-ment engineering telemetry

• Receiving, archiving, and reducing all instru-ment science data to Planetary Data System(PDS) Level 1 format

• Distributing science data and associated en-gineering/navigation data to co-investigators,participating scientists, the education andpublic outreach team, and the PDS

• Receiving and archiving higher-level sciencedata products produced by the science team

However, for the purposes of this study, we as-sumed that the JHU/APL Mission Data Center(MDC) will prepare Level 1 telemetry and dis-tribute it to the Science Team Operations Cen-ter (STOC) and the Payload Operations Centers(POCs). Planning for science operations will useapplications built on the JHU/APL Science Plan-ning Framework and Toolbox. Clock-correlationprocessing is performed by the MOC and dis-seminated to the STOC, the POCs, and the navi-gation team.

JPL’s Navigation Group provides radiometric dataconditioning and validation, Doppler data pre-con-ditioning, orbit determination, ancillary navigationdata processing, and verification of JHU/APL-computed maneuvers. It interfaces directly withthe DSMS for tracking and acquisition data.



4.6.3 Summary

The ground and data system architecture identi-fied will meet currently identified requirements;it also takes of advantage of significant heritagefrom previous programs. Once the instrumentteams have been identified, this architecture canbe refined to suit both these teams’ specific needsand those of the Solar Probe mission.

4.7 Mission Operations

4.7.1 Requirements

The Solar Probe Mission Operations System(MOS) consists of the teams and ground facili-ties required to conduct post-launch operations.It must to support the operational concept de-fined in Section 4.2.

4.7.2 Approach

Because of the similarities in their missions,Solar Probe builds on the successful experiencesand lessons learned from NEAR, CONTOUR,and MESSENGER.

The MOC houses the elements used to conductthe three main functions within space opera-tions—activity planning and scheduling, real-time command and control, and off-lineperformance assessment. The Solar Probe MOTwill take advantage of lessons learned on pre-vious missions and build on the successful ex-periences from NEAR, CONTOUR, andMESSENGER by capitalizing on existing in-frastructure and utilizing the same proven pro-cesses and procedures used to conduct thepost-launch operations. In addition, software toperform these various functions will be exten-sively reused, reducing the cost and risk of ad-ditional software development. The MOT isalso supported by outside organizations in spe-cific areas. As they have done on previous mis-sions, JPL will provide navigation support. Theinstrument suite POCs will provide the instru-ment command sequences to meet the detailed

science objectives as defined by the scienceteam.

In operations planning, the design, development,test, review, and control of the mission-criticalsequences such as science acquisition activitiesand TCMs adhere to the proven process em-ployed on past missions. Before any sequenceis uplinked, it is run both on the real-time hard-ware-in-the-loop simulator and on the faster-than-real-time software simulator.

During real-time contact with the spacecaftthrough the DSMS, routine and nonroutine op-erations are performed to ensure the health andsafety of the spacecraft. During such contacts,the telemetry is evaluated for out-of-limit con-ditions and to verify the spacecraft’s overall stateof health. Commands are uploaded to the space-craft for housekeeping functions and for sciencedata acquisition sequences. Also performed dur-ing real-time contacts on a less frequent basisare parameter uploads, flight software uploads,and TCMs as required. In the event of non-nomi-nal spacecraft behavior, the MOT implementsdiagnostic procedures and initiates other pre-scribed responses where applicable. A summaryof the types of activities performed duringDSMS contacts in the various mission phasesin shown in Table 4-14.

Off-line performance assessment is performedon the spacecraft as well as on the overall op-eration itself to continually evaluate the effi-ciency of the operation and to incorporatelessons learned throughout the mission. Assess-ment of the spacecraft includes both short- andlong-term trending and the off-line resolutionof anomalous spacecraft behavior. In this area,the engineering development teams remain onthe project at a low level to evaluate their sub-systems’ performance throughout the mission.

The science team defines the science require-ments and the science acquisition activities ona monthly basis. From this, the MOT and



Figure 4-33. Timeline of a 6-month window around a solar encounter.

P-90d P-60d P-30d P-10d P+10d P

P+30d P+60d P+90d

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02-0817R-35

Table 4-14. Mission Operations Team activities during mission phases.

Mission Phase Duration Contact Fre-quency

Activities Performed

Spacecraft separationand early checkout

1 week 24 h/day • Performance monitoring• Initial turn-on and checkout of spacecraft subsys-

temsContinued spacecraftcheckout

12 weeks 3 or 4 to 8 h/week • Continued performance monitoring• Initial instrument turn-on and checkout• Further checkout of spacecraft functionality and

performance• Spacecraft certification

Cruise 1 10 months 1 to 8 h/week • State-of-health evaluation• Noncoherent ranging• Routine housekeeping functions• Diagnosis and initial response to anomalistic

behavior• Collection and downlink of cruise science (approx.

22 Mbits/week)Jupiter flyby 2 months 3 or 4 to 8 h/week • Normal cruise functions

• Plan, test, implement, analyze navigation burnCruise 2 20 months 1 to 8 h/week • Normal cruise functionsSolar encounter 1 preps 12 weeks 3 or 4 to 8 h/week • Normal cruise functions

• Plan, test, implement, and analyze required TCMs.• Rehearsals of encounter 1 sequence on

simulators• Rehearsals of encounter 1 sequence on space-

craft• Updates to sequence based on tests• Additional rehearsals on simulators and spacecraft

Solar encounter 1 20 days 24 h/day • Normal cruise functions• Maintain continuous downlink• Change spacecraft configuration as required• Change downlink rates as required• Update and retest upcoming sequences as

situations ariseEncounter 1 data retrieval 8 weeks 2 to 8 h/day • Normal cruise functions

• Plan, test, implement, and analyze re-targetingburn

• Playback SSR data from both SSRsCruise 3 46 months 1 to 8 h/week • Normal cruise functionsSolar encounter 2 preps 12 weeks 3 or 4 to 8 h/week • Same as encounter 1Solar encounter 2 20 days 24 h/day • Same as encounter 1Encounter 2 data retrieval 8 weeks 2 to 8 h/day • Same as encounter 1

instrument suite POCs derive the requiredspacecraft and instrument configurations. Theinstrument suite POCs construct the requiredinstrument command sequences to perform thescience acquisition process fortheir respective instrumentsuites. The MOT constructsthe spacecraft command se-quences for attitude maneu-vers, data recording, and anyother required spacecraft con-figuration changes. Prior toand during the flyby encoun-ter, the planning system is de-signed to allow late changes in

command sequences as situations arise, pro-vided there is ample time for test. A 6-monthtimeline with the solar encounter in the middleof the window (Figure 4-33) shows the



types of activities performed by the MOT andspacecraft.

The mission design and navigation teams par-ticipate with the MOT in maneuver definitionand implementation. In addition, the navigationteam processes the noncoherent ranging infor-mation to compute the spacecraft ephemeris.

As has been the case on prior missions, the MOTstaff will increase in size to support the solarencounters and maneuvers and decrease to asmall core team during cruise phases. Althoughthe team members will be shared with other on-going missions, because of the team’s waningand waxing nature and because the Solar Probemission is so long, it is critical that a long-termknowledge management and retention plan beimplemented.

To support this important effort, a training pro-gram will stress the roles and responsibilities ofeach member of the MOT. This not only expe-dites new member training, but also assists in theadvancement or cross-training of existingmembers. The training program will consist of acombination of videotaped classroom lecture ses-sions and “practical lab” training on the groundsimulator to reinforce the classroom training. Thetraining program will conclude with a certifica-tion process that must be completed by all MOTmembers before they are placed into the on-linerotation without assistance.

During the development phase of the mission, theMOT develops the MOS in addition to gaining a

detailed understanding of the operation of thespacecraft. Members of the MOT support the I&Teffort, and members of the I&T team support theMOS development through assistance with se-quence development and test. In the I&T phase,the MOT is allocated spacecraft test time duringwhich mission simulations are conducted in a “testit as you fly it” approach. These mission simula-tions are conducted with three main objectives inmind. First, they test the capability of the MOSto construct the sequences, load them, and verifyexecution. Second, the tests provide an excellentfull-up spacecraft test where all components ex-cept the power system are exercised as they wouldbe post-launch; the power system will be simu-lated, of course. Third, the testing provides an ex-cellent opportunity for MOT training with thespacecraft. The mission simulations consist of alaunch, separation, and early on-orbit checkoutsimulation, a typical TCM, the Jupiter flybysequence, and a solar encounter simulation. Thissolar encounter sequence developed prior tolaunch is refined after post-launch calibrations.The mission simulation tests are archived for fu-ture training opportunities.

4.7.3 Summary

The mission operations approach supports themission requirements and current operationalconcept defined in Section 4.2. A key elementof the MOS is the emphasis on knowledgemanagement and retention so that cost-driven variations in staffing do not impactperformance.


5: Risk Mitigation

5. RISK MITIGATION

This report has described in detail an optimizedengineering solution for a viable Solar Probemission. The mission design selected maximizesscience return by affording multiple quadraturepasses of the Sun within one solar cycle. Thehighly fault-tolerant flight system carries a pay-load of in situ and remote sensing instrumentsinto the harsh environment of the Sun’s corona.Science data are both telemetered in real timeand redundantly recorded onboard for laterdownlink, ensuring maximum data return.

As the mission and spacecraft designs developed,trade studies were performed to examine the pos-sible means of achieving the mission objectivesand establishing the requirements for each sub-system. In addition to technical and cost consid-erations, balancing overall risk was a criticalelement of all trade study evaluations. To arriveat an optimal solution, we included technical risk,cost risk, and improved probability of missionsuccess as discriminators in the decision process.The result is a feasible design that requires mod-est technology development only where that de-velopment is uniquely necessary to achieve theprimary mission objectives. Table 5-1 presentsthe heritage and maturity—as reflected by thetechnology readiness level (TRL)—for the SolarProbe subsystem components.

The balanced risk approach identified severalkey technical challenges for Solar Probe. Tomitigate these technical challenges, a risk retire-ment program has been outlined, beginning 2years prior to mission formulation. Such a sched-ule allows time for the iteration and revision ofprocesses that are inevitable in technology de-velopment activities.

For each required technology development area,a roadmap has been developed that identifies thesteps to be taken to bring the system to matu-rity. In each case, contingency approaches have

been identified should the planned developmentbe unsuccessful. The impact of having to imple-ment fallbacks is also identified. Appropriatetrigger dates and technical criteria will be iden-tified that specify at what point each develop-ment effort must prove successful or the fallbackoption will be initiated.

From the list in Table 5-1, the systems requiringtechnology development to advance their readi-ness level have been identified as those with aTRL rating below 5. These include thermal/structure, telecommunications, dust protection,and attitude control. While the power system alsorequires maturation, this development is largelyout of the scope of the Solar Probe missionimplementers, and is being directed by the De-partment of Energy (DOE) and the NASANuclear Systems Initiative.

The remainder of this section addresses the tech-nical areas identified as requiring technologydevelopment. The nature of the risk is described,followed by a summary of the relevant state ofthe art of the technology. The planned steps to-ward mitigation are laid out, along with afallback plan and the impact of that contingencyaction.

5.1 Thermal Protection System (TPS)

The system requiring the most technology de-velopment, and whose fallback position has thegreatest impact on Solar Probe, is the ThermalProtection System (TPS), which includes theprimary heat shield assembly, the secondaryshield assembly, and light tubes that penetratethese shields to allow the nadir-pointing remotesensing instruments to see the Sun.

5.1.1 Primary Shield Assembly

A large (2.7-m diameter, 5-m height) primaryshield is required to shadow the spacecraft


5: Risk Mitigation

Table 5-1. Subsystem components, heritage, and technology readiness levels (TRLs).

Subsystem Vendor Heritage TRL CommentsIntegrated Electronics Module

DC/DC converter JHU/APL Modified CONTOUR 7 Modifications for higher efficiencySolid state recorder JHU/APL New development 6 New impl. of existing technologyFlight computer BAE First flight with Deep Impact 6 Off-the-shelf procurementCmd/TLM card JHU/APL New development 6 New impl. of existing technologyDownlink card JHU/APL Modified CONTOUR 7 Frequency and control pass-through to modsUplink card JHU/APL Modified CONTOUR 7 Minor mods, customized for Solar Probe

Attitude Interface UnitProcessor board BAE First flight with Deep Impact 6 Off-the-shelf procurementAttitude interfaceelectronics

JHU/APL Slight mod of several JHU/APL builtflight units 7 Minor mods., customized for Solar Probe

RF CommunicationsUSO JHU/APL Cassini 9 Build to printX-band SSPA JHU/APL Modified MESSENGER 7 Minor mods, customized for Solar ProbeX to Ka freqconverter

JHU/APL New development 2 Migrate to Ka-band frequencies

Ka-band SSPA JHU/APL New development 2 Develop high efficiency 8W RF outputDiplexer MCC CONTOUR 9 Off-the-shelf procurementRF switch assy Com-Dev CONTOUR 9 Off-the-shelf procurementMGA JHU/APL Modified CONTOUR 7 Minor mods, customized for Solar ProbeHGA JHU/APL New development 2 Design for both Ka & X-band operationLGA JHU/APL Modified CONTOUR 7 Minor mods, customized for Solar Probe

Attitude Control SystemStar trackers Sodern 20 flight units deli; 3 in flight 9 Off-the-shelf procurementIMU Northrop-

GrummanUpdated NEAR design; will haveflown on MESSENGER 7 Off-the-shelf procurement

DSADs Adcole Many flight programs 9Solar horizon sensor JHU/APL New development 3 Current concept based on µDSAD chip

PowerPower distr. unit:

Power switching JHU/APL Mod of MESSENGER, STEREO 7 Minor mods, customized for Solar ProbeCmd decoder JHU/APL Mod of MESSENGER, STEREO 7 Minor mods, customized for Solar Probe1553 Board JHU/APL Mod of MESSENGER, STEREO 7 Minor mods, customized for Solar Probe

Shunt regulator unit:Shunt regulator JHU/APL Mod of NEAR, CONTOUR, New

Horizons 7 Minor mods, customized for Solar Probe

Capacitor bank JHU/APL New Horizons 7 Mod. capacitance and small mods to electronicsLithium ion battery Eagle-Picher New development 7 New implementation of existing technologyBoost converter JHU/APL Modified ACE 7 Customized for Solar ProbeCharge controller JHU/APL Based on several JHU/APL flight

programs 7 Minor mods, customized for Solar Probe

RTG GFE New development 7 MMRTG being developed by DOE. Technologysame as previous RTGs

PropulsionTank Gen

DynamicsEURECA 9 Off-the-shelf procurement

Thrusters Gen.Dynamics

Many flight programs 9 Off-the-shelf procurement

Valves, filters VACCO Many flight programs 9 Off-the-shelf procurementPressuretransducers

Paine Many flight programs 9 Off-the-shelf procurement

MechanicalTelescoping boom Astro Aerosp. Modification of existing designs 7 Modified for Solar Probe applicationRotary actuators Moog Many flight programs 9 Off-the-shelf procurementDust shield JHU/APL New development. 3 Potential use of CONTOUR materials & design

approachThermal

Primary TPS JHU/APL New development 2 Large carbon–carbon structures have been built forspace applications (e.g., shuttle)

Secondary TPS JHU/APL New development 2 Need to select insulation materialLight tubes JHU/APL New development 2


5: Risk Mitigation

systems and science payload from solar radia-tion while minimizing mass loss due to subli-mation at the extreme temperatures of thenear-Sun environment. The baseline materialselected for the primary heat shield is carbon-carbon (C-C), which is chemically and physi-cally stable in the deep-space environment, isinsensitive to hydrogen embrittlement, and hasa high strength-to- weight ratio. It is also usedfor the support struts and other assembly com-ponents. The specific risks associated with theprimary shield assembly are (1) itsmanufacturability and (2) its outgassing perfor-mance at extreme temperatures.

5.1.1.1 Manufacturability

Because of its size and mass, the primary shieldrequires a robust system of struts to support itand connect it with the spacecraft bus. The pri-mary shield assembly is a critical component ofSolar Probe, and its manufacturing and assem-bly processes should be demonstrated as earlyas possible to ensure that they are mature whenmission implementation begins.

State of the Art. Large carbon-carbon structuresnot unlike that proposed for Solar Probe are com-monly used for high-temperature applications,such as the Space Shuttle nose cones. Thesestructures are about 2 m in diameter and are sub-jected to temperatures of approximately 2000K. The fabrication processes of such structuresare largely the same as those required for theSolar Probe primary shield.

The fundamental shield design (a 15° half-anglecarbon-carbon cone with a diameter of about 2.5m) has been studied in depth during previousSolar Probe studies, and a manufacturer of car-bon-carbon materials has been involved in thecurrent study specifically to address the feasi-bility of fabricating the primary shield assem-bly. A methodology evaluation has beenperformed and several alternative approaches to

assembling the shield and its strut assembly havebeen laid out.

Mitigation Plan. The risks associated with thefabrication of such a large assembly can mosteasily be retired by fabricating and testing full-sized prototype units. A qualified vendor willbe selected who will work with us on the devel-opment of the entire thermal protection system.A detailed approach to the fabrication of theprimary shield will be developed, including suchdecisions as whether to build multiple flat pan-els that are joined to form a faceted cylinder, orto join larger curved pieces. Such decisions willbe based on the capabilities of the selectedvendor’s facilities as well as the dimensionalrequirements of the shield.

In parallel with this development, materials test-ing will be performed to verify the mechanicalproperties of the samples. Tests will be per-formed on sample joints and interfaces betweenthe various primary shield assembly compo-nents. The results of these tests will be used todevelop mature assembly procedures. The cur-rent plan is to fabricate two full-scale prototypeshield assemblies in series. This conservativeapproach allows lessons learned during the firstmanufacturing run to be implemented andproven in the second run.

Fallback and Implications. The true risk asso-ciated with the fabrication of the shield assem-bly is not whether the assembly can be made,but whether the mass estimated for a stiff andsurvivable structure is adequate. Going to thetrouble of building prototypes will not only re-tire manufacturing risk, but also reduce the un-certainty associated with the mass allocation ofa major component, and should improve theoverall mass margin. If complications arise dur-ing the manufacturing process that are insur-mountable within the allotted schedule andbudget, then the mass allocation to the shieldwould be increased at the expense of margin.


5: Risk Mitigation

5.1.1.2 High-Temperature Behavior

Although it is well known that carbon-carbonmaterials can withstand the extreme tempera-tures that Solar Probe will encounter, the actualtemperature that the shield reaches will directlyaffect the extent of outgassing that occurs. Aspecific mass sublimation limit has been deter-mined, above which the primary science objec-tives are jeopardized (see discussion in Section4.4.2.1).

State of the Art. Candidate carbon–carbon ma-terials and manufacturing processes have beenextensively evaluated in previous Solar Probestudies, such as that conducted by LockheedMartin, JPL, and NASA Langley Research Cen-ter (Dirling 1998). These studies measured op-tical properties (in particular absorptivity � andemissivity �) of carbon–carbon at high tempera-tures. Mass loss at high temperatures has alsobeen characterized, but the data are sparse, andthe tests were limited to small samples and near-normal incidence angles.

The very limited amount of high-temperaturetest data with these types of materials createsdesign uncertainty despite significant margin inboth design temperature and solar incidenceangle.

Mitigation Plan. A detailed plan for high-tem-perature testing of carbon–carbon materialsamples must be developed that includes bothoff-angle optical property measurements andmass loss measurements under the same condi-tions. Solar furnaces such as that at Sandia Na-tional Laboratories are candidate facilities forsuch tests. Obtaining results consistent with pre-vious studies using targeted samples of carbon–carbon material will provide the neededconfidence in the previous experience and de-sign margin.

Fallback and Implications. The 2-year pre-for-mulation risk reduction period allows time for

repeated test series using a variety carbon–car-bon samples. If the test results demonstrate thatthe predicted margin is not representative, thena modification of the shield design will be con-sidered. The primary option would be to de-crease the cone angle below 15° to reduce theshield temperature. This would elongate theshield assembly and likely increase its mass.

5.1.2 Secondary Shield Assembly

While the primary shield shadows the spacecraftbus, the secondary shield provides the bulk ofthe thermal insulation between the hot primaryshield and the spacecraft bus. The temperaturegradient across this assembly will range fromthe primary shield temperature of approximately2200 K down to about 750 K at the base of thesecondary shield. As a result, this shield mustincorporate low-density materials that can with-stand extreme temperatures and exhibit low ther-mal conductivity. The specific risk associatedwith this system is the structural integrity of thecandidate materials and a mechanical assemblythat allows them to be mounted to the primaryshield.

State of the Art. High-performance materialssuch as carbon aerogel, carbon fiber batts, andaerogel-infiltrated foams offer great promise asthe high-temperature components of the second-ary shield. Thermal conductivity of these mate-rials at high temperatures is low enough toprovide efficient thermal isolation, and the lowdensities help keep the mass allocations down.Thermal analysis of such a secondary shield in-corporated into the current Solar Probe designhas confirmed the desired performance. Opticaland mass loss properties of these materials haveyet to be characterized at the extreme tempera-tures anticipated for Solar Probe, and little hasbeen done with respect to packaging into largerassemblies.

Mitigation Plan. A carefully planned strategyof testing must be developed to characterize both


5: Risk Mitigation

high-temperature and mechanical properties ofcandidate materials. The selection and testingof adhesives or other elements of the assemblymust be included as well, as the results of thesetests will have direct bearing on the assemblydesign. Full-scale prototype assemblies will befabricated and tested once the materials andpackaging methods have been determined. Me-chanical tests of such large structures arestraightforward, but high-temperature testsmight not be feasible. It is important, therefore,that reliable relationships between materialsample data at extreme temperatures and lower,more easily attainable temperatures can be made.With these, the response of full-scale prototypesat the extreme can be validated.

Fallback and Implications. If, after exploringthe candidate materials and assembly ap-proaches, it is improbable that the high-perfor-mance materials listed above can be incorporatedinto the secondary shield, lower-performance,commercial materials must be considered. Uti-lizing such materials in the shield assembly af-fords less of a challenge, but the resulting shieldwill be more massive.

5.1.3 Light Tubes

Two of the remote sensing instruments in thescience payload require nadir-pointing fields ofview (FOVs). For this study, we assumed thatnadir-pointing FOVs would be achieved by in-corporating carbon–carbon light tubes to createan optical path through the primary and second-ary shields. The specific risk associated with thelight tubes is the degree to which they will limitthe performance of the nadir-viewing imagersor complicate their development and integration.A related e concern is that the alignment of thetubes within the TPS shields cannot be main-tained to the satisfaction of the imagers.

State of the Art. Thermal analysis was per-formed on a variety of tube designs and it was

determined that thermal throughput could belimited to about 25 W if the tubes were trun-cated at the bottom (spacecraft side) of the sec-ondary shield. Heat rejection remained aproblem if the tubes extended all the way to theinstrument apertures.

Mitigation Plan. Before effort is expended ondealing with fabrication issues, the general con-cept of imaging through truncated, tapered tubesmust be validated through optical tests with rep-resentative instrumentation. Once this approachis validated, the fabrication and mechanical in-tegration issues need to be addressed. A fabri-cation and test plan will be developed andfollowed to produce prototype tubes and meansof attaching them to the primary and secondaryshields. Initial mechanical and alignment testswill be conducted with a mockup of the TPSshields, and once primary and secondary shieldprototypes have been built, tests can begin withthe entire assembly. In the event that the use oftruncated light tubes is not viable with the im-agers, extending them to the instrument aper-ture must be pursued, although this adds aformidable thermal control challenge for theinstrumenters.

Fallback and Implications. As with the othercomponents of the TPS, mass is at risk. Strength-ening the tubes and their supports, which willrequire increased mass, can mitigate alignmentproblems. If the truncated tubes cannot providethe imagers the unobscured view they require,then further strengthening might be required tomate the instruments to the tubes.

5.2 Telecommunications

The Solar Probe telecommunications system isdual frequency, employing X-band (for uplinkand downlink) and Ka-band (for downlink only).The addition of Ka-band capability allows So-lar Probe to overcome the transmission lossesassociated with coronal scintillation and thereby


5: Risk Mitigation

to support real-time downlink of science dataduring solar passes. Given the 20-W power al-location for the telecommunications system andthe predicted performance of the 0.8-m dual-fre-quency high gain antenna (HGA), a relativelyhealthy solar encounter data rate of ≥25 kbpscan easily be achieved with a high-efficiency(approximately 40%) solid-state power ampli-fier (SSPA). The specific risk associated withthis system is the successful development of sucha high-efficiency SSPA.

State of the Art. Currently available Ka-bandSSPAs achieve efficiencies of only about 20%,whereas 40% efficient X-band SSPAs will soonbe obtainable. Ka-band antennas are readilyavailable with efficiencies of 60%, the highestpublished efficiency currently being 73%. MarsGlobal Surveyor incorporated a dual frequencyX- and Ka-band antenna, but that system wasnot optimized for efficiency.

Mitigation Plan. The efficiencies of the SSPA,the antenna, and the electronics packaging allcontribute to the overall performance of the Ka-band system. The 40% efficient SSPA require-ment assumes only average performance from theantenna and packaging. An increase in antennaperformance relaxes the efficiency requirementfor the SSPA, so improvements in that develop-ment will also be pursued. It is currently conser-vatively assumed that the dual-frequency antennacan achieve only 50% efficiency, although betterperformance is likely. If an antenna efficiency of70% can be attained, then the required Ka-bandSSPA needs to be only about 25% efficient toafford the desired 25 kbps data rate.

The development process for the Ka-band sys-tem begins with the down-selection of a pre-ferred technology, such as gallium arsenide(GaAs), gallium nitride (GaN), indium phos-phide (InP), or silicon carbide (SiC). Once thepreferred technology is chosen, parts are selectedthat best suit the technology, and the fabrication

and test of a prototype amplifier can ensue. Inparallel, a prototype dual-frequency antenna willbe developed and its performance characterized.As with the other technology development ef-forts, time is allotted to accommodate redesignsand repeated tests in both of these efforts.

Fallback and Implications. In general, the lowerthe peak telecommunications system efficiencyachieved, the lower the data rate available dur-ing solar encounters. As a result, the fallbackposition is to accept a lower data rate at perihe-lion. If the SSPA development only accom-plishes 25% efficiency and the dual-frequencyantenna cannot be optimized above 60%, thenthe resulting data rate will be 75% of the de-sired value. The impact of such a fallback is thereduction of real-time data downlink, but thereis no associated science loss, since the real-timedownlink itself is redundant. All science datacollected will be transmitted after the solar pass.

5.3 Dust Protection

A dust protection system must be incorporatedinto Solar Probe to shield the spacecraft fromparticulates of varying sizes and velocities. Spe-cific risks associated with small particles in-clude the pitting of optics, degradation of cableinsulation, and abrasion of thermal surfaces.Larger particles could cause more severe dam-age to the spacecraft, since collisions of 450km/s are possible.

State of the Art. The solar dust environment isnot well understood, due to limited data andmodels based on 1-AU observations. Analyti-cal models (hydrocodes) exist only for veloci-ties of about 30 km/s and cannot reliably beextrapolated to the 400 km/s regime.

Mitigation Plan. To clarify the current state ofunderstanding of the solar dust environment, aworkshop can be convened that will focus onthe Solar Probe environment. The purpose willbe to consolidate and interpret data, update


5: Risk Mitigation

models, and propose additional measurementsto be made.

Additionally, a model validation program needsto be developed to assess the effectiveness ofdust mitigation approaches. A national surveyof state-of-the-art equipment and facilities forlaser damage effects testing should be con-ducted. A test plan will be devised that com-bines analytical model development with dustimpact experiments to verify those models. Theeffects of dust impacts on spacecraft componentsand coatings can be assessed, and a dust protec-tion system can be evolved that incorporates adiscrete shield, localized shielding, and embed-ded packaging concepts.

Fallback and Implications. Due to the uncer-tainty of this environment, spacecraft protectionmust be maximized within the mass constraintsof the program. If analyses and models indicatea harsher environment and greater threat thancurrently assumed, then the mass margin willdecrease as more mass is allotted to dust pro-tection.

5.4 Attitude Control

To augment the star trackers and the inertialmeasurement unit (IMU), a solar horizon sen-sor is proposed as an added precaution to main-tain a Sun-pointing attitude regardless of the stateof primary or redundant attitude system sensorsduring a solar pass. The specific risk associatedwith this system is the loss of an additional safe-guard if this technology is not developed. Whilethere is ample heritage for this sensor approach,the application and environment are new.

State of the Art. Sensors performing the tasksrequired of this unit have been developed forprevious missions, e.g., the Sun gates on Viking,Voyager, and Galileo. A variety of sensor tech-nologies exist that might be applicable to thisdevelopment.

Mitigation Plan. More than one design conceptfor this sensor will be developed. One approachis to detect the solar horizon optically, usingvariations of digital solar attitude sensors or simi-lar devices. A method that does not rely on im-aging the horizon directly incorporatesthermistors located around the base of the pri-mary shield that indicate when the shield be-gins to point away from Sun center. Prototypesof these sensors will be fabricated and validatedthrough representative tests.

Fallback and Implications. If a complex sys-tem incorporating digital Sun sensors cannot bedeveloped for moderate cost, the simpler ther-mistor system should be utilized. If this approachcannot be validated, the risk of relying on theredundant attitude control system devices mustsimply be accepted.

5.5 Radioisotope Power System

Using a radioisotope power system enables So-lar Probe to operate reliably throughout its mis-sion life and at distances where solar power isnot feasible (farther than 2 AU and closer thanabout 0.3 AU). Due to the extremely limitedavailability of traditional radioisotope thermo-electric generator (RTGs), Solar Probe will uti-lize the new, smaller multi-mission radioisotopethermoelectric generators (MMRTGs) that arebeing developed by NASA and the Departmentof Energy (DOE) for a wide range of future spacemissions. The performance of general-purposeheat source (GPHS) RTGs and issues related totheir safety are well understood and pose littletechnical risk to Solar Probe. The specific risksto the mission are programmatic in nature, andconcern the timely development of theMMRTGs and completion of the launch ap-proval process.

Current Status. RTGs have been used on morethan 15 NASA missions over the past 30 years,and the design has remained unchanged since the


5: Risk Mitigation

Galileo mission. The MMRTG currently beingdeveloped is a modification of this design thatuses a shorter housing holding only eight GPHSmodules. The MMRTG development is part ofNASA’s Nuclear Systems Initiative, and is beingcoordinated with the DOE. A competitive pro-curement has been initiated for the design, de-velopment, and qualification of MMRTGs, andthe development schedule calls for the first flightunit to be available in early 2008.

Mitigation Activities. Representatives of theSolar Probe team will work closely with DOE,NASA, and the MMRTG developer to ensurethat mission requirements are adequately repre-sented and that the details of the spacecraft de-sign remain fully compatible with the design ofthe MMRTG. The overall development sched-ule will also be closely monitored.

In addition, the significant effort that is antici-pated to meet all safety requirements typicallyassociated with RTG missions will begin in thepre-formulation phase. Solar Probe will have thegreat benefit of leverage from the experience ofthe New Horizons mission team, which has beenactively engaged in these activities. An early start

is the best mitigation against delays associatedwith the documentation preparation and approvalprocesses.

Implications. The Solar Probe mission conceptis based on a strawman launch date in May 2010,which is well matched to the current MMRTGdevelopment schedule. The launch opportunityrecurs every 13 months with only minor changesto the mission profile, but the later in the mis-sion development that such a slip occurs, themore costly that delay becomes. Early and closeparticipation in the MMRTG development ef-fort can help reduce such risk.

5.6 Risk Mitigation Activity Schedule

Figure 5-1 shows a high-level timeline of SolarProbe risk mitigation activities. Phase A andPhase B are consistent with the strawman mis-sion launch date of May 2010, and are includedhere simply to illustrate that risk reduction ac-tivities continue into mission implementation.Two years of technology development prior toPhase A is considered sufficient to perform thetasks outlined above and to achieve the goal ofretiring risk early.

Primary Shield

Pre-Phase A Phase A Phase B

Dust Shield

Secondary Shield

Light Tube

Ka-Band SSPA

FY03 FY04 FY05 FY06

Material Testing FabricationAssy.

TestingNEEDDATE

Material Testing Fabrication Assy.Testing

NEEDDATE

Material Testing Fabrication Assy.Testing

Perf.Testing

TechnologyDownselect

PartsDownsel.

Prototype Board Development NEEDDATE

Workshop AnalyticalModeling

Dust ShieldDesign

NEEDDATE

LaserTest

LaserTest

ResultsAnalysis

Figure 5-1. Risk mitigation schedule.02-0817R-47

Solar Probe: An Engineering Study A-1

APPENDIX A

SOLAR PROBE MASS, POWER, AND SPACECRAFT DIMENSIONS

Subsystem Name Qty.

CBE Mass Each (kg)

CBE Mass Total (kg)

Growth Allowance (Reserves)

Total Mass Not to

Exceed (incl/Growth Allowance)

Notes and Heritage

39.70 54.58

Remote Sensing, Data Unit 1 3.00 3.00 40% 4.20 Drawn from AO responses, 1997 NRARemote Sensing,Power Unit 1 3.00 3.00 40% 4.20 Drawn from AO responses, 1997 NRAEUVI (Requires Light Tube) 1 3.00 3.00 40% 4.20 Drawn from AO responses, 1997 NRAEUVI Light Tube 1 0.50 0.50 40% 0.70 APL Structures EstimateASC 1 2.80 2.80 40% 3.92 Drawn from AO responses, 1997 NRAVMH (requires Light Tube) 1 3.00 3.00 40% 4.20 Drawn from AO responses, 1997 NRAVMH Light Tube 1 0.50 0.50 40% 0.70 APL Structures Estimate

Insitu, Data Unit 1 3.00 3.00 40% 4.20 Drawn from AO responses, 1997 NRAInsitu, Power Unit 1 4.50 4.50 40% 6.30 Drawn from AO responses, 1997 NRAPWS, DPU and Search Coils 1 3.50 3.50 40% 4.90 Drawn from AO responses, 1997 NRAEPCS, EPD Sensor head and electronics 1 0.70 0.70 40% 0.98 Drawn from AO responses, 1997 NRAFSWD, Electronics and Sensor Head 1 1.00 1.00 40% 1.40 Drawn from AO responses, 1997 NRASWICES Instrument Total Allocated 1 4.40 4.40 40% 6.16 Drawn from AO responses, 1997 NRAMAG, DPU and Sensor 1 1.80 1.80 40% 2.52 Drawn from AO responses, 1997 NRAScience Boom 1 5.00 5.00 20% 6.00 Astro-Aerospace telescoping composite boom, 5 m long

16.920 20.30High Gain Antenna 1 4.00 4.00 20% 4.80 APL RF System Estimate, .8 m antennaHGA Antenna Drive Motor 2 1.30 2.60 20% 3.12 Moog Rotary Actuator Type 2HGA Mechanical Support Structure 1 1.50 1.50 20% 1.80 APL Structures EstimateHGA Motor Electronics 1 1.00 1.00 20% 1.20 Stereo HeritageUltra Stable Oscillators 2 1.50 3.00 20% 3.60 New Horizons DesignLow Gain Antennas 2 0.35 0.70 20% 0.84 NEAR, ACE, Contour HeritageMedium Gain Antenna 1 0.56 0.56 20% 0.67 APL RF System EstimateX-Distrubution 1 0.80 0.80 20% 0.96 APL RF System EstimateCables, etc. 1 2.76 2.76 20% 3.31 APL RF System Estimate

40.58 48.24IMU 1 6.60 6.60 10% 7.26 LittonReaction Wheel 4 5.70 22.80 20% 27.36 ITHACH-TW-4A120Reaction Wheel Bracket 4 0.50 2.00 20% 2.40 TIMED Heritage, scaledStar Cameras 2 3.01 6.02 20% 7.22 SODERN, SED-16DSAD’s 2 0.26 0.52 20% 0.62 Stereo DesignDSAD Electronics 1 0.64 0.64 20% 0.77 Stereo DesignHorizon Sensors 1 1.00 1.00 40% 1.40 APL System Engineering EstimateStar Camera Bracket 1 1.00 1.00 20% 1.20 APL Structures Estimate

95.80 114.96

RTG3 24.00 72.00 20% 86.40

Multi-Mission Module Baseline, BEST case mass, Emg, 5/23/02

Shunts 2 0.50 1.00 20% 1.20 Based on New Horizons, 282 watts BOLSRU (Capacitor Bank) 1 6.30 6.30 20% 7.56 New Horizons, mod based on Uno Carlsson inputsBattery Charger 1 0.50 0.50 20% 0.60 Uno Carlsson EstimateBattery Boost Converter 1 1.00 1.00 20% 1.20 Uno Carlsson EstimateBattery(s) 1 2.50 2.50 20% 3.00 SAFT V34570 D, 4.6 Ahr. Case estimate DEPDU/AIU 1 12.50 12.50 20% 15.00 New Horizons, based on 282 watts, 1 RTG.

CBE Instrument Mass Total

Table A-1. Solar Probe Equipment List and Detailed Mass Breakdown

Telecommunications

Instruments

Power System,

Remote Sensing Instruments

Insitu Instruments

Tellecommunications System Current Total

Guidance and Control (G&C) System,CBE G&C System Current Total=

CBE Power System Total=

A-2 The Johns Hopkins University Applied Physics Laboratory

Subsystem Name Qty.

CBE Mass Each (kg)

CBE Mass Total (kg)

Growth Allowance (Reserves)

Total Mass Not to

Exceed (incl/Growth Allowance)

Notes and Heritage

Table A-1. Solar Probe Equipment List and Detailed Mass Breakdown

148.80 190.96Primary Heat Shield 1 39.00 39.00 20% 46.80 APL Structures Estimate, 2.8 meter diameterPrimary Heat Shield Support System 1 10.90 10.90 20% 13.08 APL Structures EstimateSecondary Heat Shield Aerogel 1 62.00 62.00 40% 86.80 APL Structures Estimate, 19 cm .06 gm/cm^3Secondary Heat Shield face sheets 3 5.70 17.10 20% 20.52 APL Structures Estimate,SC MLI 1 18.90 18.90 20% 22.68

, ( ,in**2 area)

Heaters 1 0.00 0.00 20% 0.00 APL Thermal EstimateDiode Heat Pipe 1 0.90 0.90 20% 1.08 Messenger Design

14.41 17.30Integrated Electronics Module (IEM) 2 6.85 13.70 20% 16.44 Contour/Messenger/STEREO HeritageTRIOS 14 0.05 0.71 20% 0.86 Contour/Messenger/STEREO Heritage

21.83 24.42Tank 1 10.89 10.89 10% 11.98 PSI P/N 80409-1, Centaur Upper Stage HeritagePiping 1 2.27 2.27 20% 2.72 APL Propulsion System Estimate / Pluto HeritageThrusters, 1 lb 16 0.36 5.79 10% 6.37

j ( ) g y, gHeritage

Filter 1 0.16 0.16 10% 0.18 Vacco P/N FOD10635, CONTOUR/Pluto Heritagelatch Valves 2 0.34 0.68 10% 0.75 Vacco P/N V1E10747, CONTOUR/Pluto HeritagePressure Transducer 1 0.22 0.22 20% 0.26 Paine P/N 213-76-260-02, CONTOUR/Pluto HeritageElectrical Connectors 10 0.02 0.23 20% 0.27 APL Propulsion System EstimateCabling 1 1.36 1.36 20% 1.63 APL Propulsion System EstimateFill and Drain Valves 2 0.11 0.23 10% 0.25 Vacco P/N V1E10433, CONTOUR/Pluto Heritage

Primary Structure 85.60 20% 102.72 15% of CBE wet mass allocatedDust Shield 12.00 20% 14.40 Eng/Willey Allocation, matches with JPL MemoHarness 21.55 20% 25.86 Allocation based on Timed/Contour/Messenger heritage

Best Estimate SC Dry Mass 497.19 613.73 Useable Propellant 69.10 69.10 Residual Propellant 1.37 1.37 Pressurant 0.60 0.60Total consumables 71.07 0% 71.07CBE SC Mass Wet 568.26 684.80CBE Max Dry Mass 641.93 641.93Max SC Launch Mass 713.00 713.00Mass Margin Wet 25.47% 4.12%Mass Margin Dry 29.11% 4.59%Unalocated Reserves (Kg.) 144.74 28.20

Sub Totals

C&DHCBE C&DH System Total=

PropulsionCBE Propulsion System Total=

CBE Thermal Sub System Total=Thermal

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Solar Probe: An Engineering Study B-1


APPENDIX B

REFERENCES

Armstrong, J. W., and R. Woo, Contribution toStarprobe Report, Inter-Office Memorandum no.3331-80-070, Jet Propulsion Laboratory, Jet Pro-pulsion Laboratory, Pasadena, CA, Dec. 15,1980.

Banjeree, D., L. Teriaca, J. G. Doyle, and P.Lemaire, Polar plumes and inter-plume regionsas observed by SUMER on SOHO. Solar Phys.,194, 43–58, 2000.

Beyer, P. E., D. J. Mudgway, and M. M.Andrews, The Galileo Mission to Jupiter: Inter-planetary Cruise Post-Earth-2 EncounterThrough Jupiter Orbit Insertion, The Telecom-munications and Data Acquisition ProgressReport 42-125, 1996, Jet Propulsion Laboratory,Pasadena, CA, pp. 1-16, May 15, 1996.

Bokulic, R. S., and W. V. Moore, Near EarthAsteroid Rendezvous (NEAR) Spacecraft So-lar Conjunction Experiment, J. Spacecraft Rock-ets, 36(1), 87–91, 1999.

DeForest, C. E., S. P. Plunkett, and M. D.Andrews, Observation of polar plumes at highsolar altitudes, Astrophys. J., 546, 569–575,2001.

Dirling, R. B. Jr., Solar Probe Thermal ShieldDesign—Carbon-Carbon Characterization,SAIC Technical Report 98/1039, Sept. 4, 1998.

Divine, N., and H. Garrett, Charged particle dis-tributions in Jupiter’s magnetosphere, J.Geophys. Res., 88, 6689, 1983.

Frazer, R. K., Solar Probe Secondary HeatShield Material Options, Technical Memoran-dum, A1C-02-116, The Johns Hopkins Univer-sity Applied Physics Laboratory, Laurel, MD,Oct. 21, 2002.

Giacobbe, F., Thermal Analysis and ConceptualDesign of the Solar Probe, Technical Memoran-dum SAI-ANYS-385, SWALES Aerospace,Inc., Beltsville, MD (in preparation), 2002.

Gloeckler, G., et al., Solar Probe: First Missionto the Nearest Star, Report of the NASA ScienceDefinition Team for the Solar Probe Mission,published for NASA by The Johns HopkinsUniversity Applied Physics Laboratory, Laurel,MD, Feb. 1999.

Jensen, J. R., and R. S. Bokulic, Experimentalverification of noncoherent Doppler tracking atthe Deep Space Network, IEEE Trans. Aero-space Electron. Sys, 36(4), Oct. 2000.

Jensen, J. R., and R. S. Bokulic, Highly accu-rate, noncoherent technique for spacecraft Dop-pler tracking, IEEE Trans. Aerospace Electron.Sys., 35(3), 963–973, July 1999.

Klaasen, K., Star Detection from Within the So-lar Corona, Jet Propulsion Laboratory, Pasa-dena, CA, Feb. 14, 2002.

Koerner, M. A., Telemetry and Command Deg-radation caused by the RF Signal AmplitudeScintillations Produced by the Near Sun PlasmaDuring the Period Near VRM Superior Conjunc-tion, Inter-Office Memorandum no. 3392-84-64,Jet Propulsion Laboratory, Pasadena, CA, May21, 1984.

Lee, S.-C., and C.-J. Lee, Solar Probe Light TubeThermal Analysis, Status Report 2002-1, pre-pared for D. Lisman under CWO#156, AppliedSciences Laboratory, Inc., Hacienda Heights,CA, June 10, 2002.

Lisman, D., Design Dust Particle Update, JPLInteroffice Memo, Solar Probe Project, June 24,2002.

B-2 The Johns Hopkins University Applied Physics Laboratory


Lisman, D., Photon Torque Analysis, ProjectAction Item #11, JPL Technical Memorandumto Dave Lehman, Solar Probe Project, Nov. 7,2001.

Mann, I., Dust Environment Near the Sun, In-stitute of Planetology, Münster, Germany, No-vember 2001 (draft report).

Morabito, D. M., S. Shambayati, S. Butman, D.Fort, and S. Finley, The 1998 Mars Global Sur-veyor Solar Corona Experiment, Telecommu-nications and Mission Operations ProgressReport TMO PR 42-142, Aug. 15, 2000.

Morabito, D. M., S. Shambayati, S. Finley, D.Fort, J. Taylor, and D. Moyd, Ka-Band and X-band observations of the solar corona acquiredduring superior conjunctions using interplan-etary spacecraft, 7th Ka Band Utilization Con-ference, Genoa, Italy, September 26–28, 2001.

Morabito, D. M., S. Shambayati, S. Finley, andD. Fort, The Cassini May 2000 solar conjunc-tion, IEEE Trans. Antennas Propagation, sub-mitted 2002.

NASA, Announcement of Opportunity: EuropaOrbiter/Pluto-Kuiper Express/Solar Probe, AO99-OSS-04, Space Science Support Office, Lan-gley Research Center, 1999a.

NASA, Solar Probe Mission and Project De-scription, Space Science Support Office, Lan-gley Research Center, 1999b.

National Academy of Sciences, The Sun to theEarth—And Beyond, A Decadal Research Strat-

egy in Solar and Space Physics, Space StudiesBoard, 2002.

Neugebauer, M., and R. Davies (eds.), A Closeupof the Sun, JPL publication 70-78, Jet Propul-sion Laboratory, Pasadena, CA, September 1,1978.

Page, D. E., Exploratory journey out of the eclip-tic plane, Science, 190, 845–850, 1975.

Sheeley, N. R., et al., Measurement of flowspeeds in the corona between 2 and 30 R

S.

Astrophys. J., 484, 47–478. 1997.

Strachan, L., R. Suleiman, A. V. Panasyuk, D.A. Biesecker, and J. L. Kohl, Empirical densi-ties, kinetic temperatures, and outflow veloci-ties in the equatorial streamer belt at solarminimum, Astrophys. J., 571, 1008–1014, 2002.

Tarditi A., et al., Photon Pressure Torque fromOff-Nadir Pointing, SAIC Technical Memoran-dum TM 01-4222-21 to D. Lisman and J.Randolph, Jet Propulsion Laboratory, Pasadena,CA, Oct. 26, 2001.

Webster, J., Magellan Spacecraft PerformanceDuring Solar Conjunction, Inter-Office Memo-randum no. 3395-94-44, Jet Propulsion Labo-ratory, Pasadena, CA, Aug. 4, 1994.

Xapsos, M. A., G. P. Summers, P. Shapiro, andE. A. Burke, New techniques for predicting so-lar proton fluences for radiation effects applica-tions, IEEE Trans. Nucl. Sci., 43(6), 2772–2777,1996.

Solar Probe: An Engineering Study C-1

Appendix C: Acronyms

APPENDIX C

ACRONYMS AND ABBREVIATIONS

� Absorptivity

� Emmissivity

µDSAD Micro Digital Solar AspectDetector

ACS Attitude Control System

AIU Attitude Interface Unit

AO Announcement of Opportunity

APL The Johns Hopkins UniversityApplied Physics Laboratory

ASCI All-Sky Coronagraph Imager

BOL Beginning of Life

bps Bits per Second

C&DH Command and Data Handling

C3

Maximum Required LaunchEnergy

C-C Carbon–Carbon

CCAFS Cape Canaveral Air ForceStation

CCD Charge-Coupled Device

CCSDS Consultative Committee forSpace Data Systems

CD Cumulative Distribution

CONTOUR Comet Nucleus Tour

COTS Commercial-Off-The-Shelf

DC Direct Current

DEIS Draft Environmental ImpactStatement

DLA Declination of LaunchAsymptote

DOE Department of Energy

DPU Data Processing Unit

DS-1 Deep Space 1

DSAD Digital Solar Aspect Detector

DSMS Deep Space Mission System(formerly the Deep SpaceNetwork, DSN)

EEPROM Electrically Erasable Program-mable Read-Only Memory

EIT Extreme Ultraviolet ImagingTelescope on SOHO

ELV Expendable Launch Vehicle

EMC Electromagnetic Compatibility

EMI Electromagnetic Interference

EPCS Energetic Particle CompositionSpectrometer

ESA European Space Agency

EUVI Extreme Ultraviolet Imager

FDR Fault Detection and Recovery

FOV Field of View

C-2 The Johns Hopkins University Applied Physics Laboratory


FPGA Field-Programmable GateArray

FSWID Fast Solar Wind Ion Detector

G&C Guidance and Control

GE General Electric

GFE Government-FurnishedEquipment

GIS Graphite Impact Shell

GPHS General Purpose Heat Sources

GSFC NASA Goddard Space FlightCenter

HGA High-Gain Antenna

I&T Integration and Test

IEM Integrated Electronics Module

IMU Inertial Measurement Unit

INSRP Interagency Nuclear SafetyReview Panel

JGA Jupiter Gravity Assist

JHU/APL The Johns Hopkins UniversityApplied Physics Laboratory

JPL Jet Propulsion Laboratory

kbps Kilobits per Second

KSC Kennedy Space Center

LASCO Large Angle and SpectrometricCoronagraph on SOHO

LGA Low-Gain Antenna

LMA Lockheed Martin Astronautics

LOS Line of Sight

LVPC Low Voltage Power Converter

LWRHU Light Weight RadioisotopeHeater Unit

LWS Living With a Star

MAG Magnetometer

MDC Mission Data Center

MESSENGER MErcury Surface, SpaceENvironment, GEochemistry,and Ranging

MGA Medium-Gain Antenna

MGS Mars Global Surveyor

MIPS Millions of Instructions perSecond

MLI Multilayer Insulation

MMRTG Multi-Mission RadioisotopeThermal Generator

MOC Mission Operations Center

MOS Mission Operations System

MOT Mission Operations Team

MPD [Solar Probe] Mission andProject Description

N/A Not Applicable

NASA National Aeronautics andSpace Administration

NEAR Near Earth AsteroidRendezvous

NEPA National EnvironmentalPolicy Act

Solar Probe: An Engineering Study C-3


NTE Not-to-Exceed

PCI Peripheral ComponentInterconnect

PDS Planetary Data System

PDU Power Distribution Unit

POC Payload Operations Center

PSAR Preliminary Safety AnalysisReport

PWS Plasma Wave Sensor

RDM Radiation Design Margin

REM Rocket Engine Module

RIO Remote Input/Output

RF Radio Frequency

RHCP Right Hand CircularPolorization

RJ

Radius of Jupiter

rpm Revolutions per Minute

RS

Solar Radius

RTG Radioisotope ThermalGenerator

S/C Spacecraft

S/N Signal-to-Noise Ratio

SDT Science Definition Team

SEC Sun-Earth Connection

SEP Sun-Earth-Probe

SIRU Space Inertial Reference Unit

SOHO Solar and HeliosphericObservatory

SPDT Single-Pole-Double-Throw

SRG Stirling RadioisotopeGenerator

SRI Southern Research Institute

SRM Solid Rocket Motor

SRU Shunt Regulator Unit

SSPA Solid-State Power Amplifier

SSR Solid-State Recorder

STEREO Solar-Terrestrial RelationsObservatory

STOC Science Team OperationsCenter

STP Solar-Terrestrial Probe

SWICES Solar Wind Ion Compositionand Electron Spectrometer

SWOOPS Solar Wind Plasma Experimenton Ulysses

TAC Thruster and Attitude Control

TCM Trajectory CorrectionManeuver

TID Total Ionizing Dose

TIMED Thermosphere, Ionosphere,Mesosphere Energetics andDynamics

TLM Telemetry

TPS Thermal Protection System

C-4 The Johns Hopkins University Applied Physics Laboratory


TRL Technology Readiness Level

TtNUS Tetra Tech NUS, Inc. (subcon-tractor to JHU/APL)

USO Ultrastable Oscillator

UVCS Ultraviolet CoronagraphSpectrometer on SOHO

VEEGA Venus-Earth-Earth GravityAssist

VMH Visible Magnetograph–Helioseismograph

VVEJGA Venus-Venus-Earth-JupiterGravity Assist

XFER Transfer

an engineering study - parker solar probeof a space mission, solar probe, to travel to the sun’s...

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