,', 

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Project Cerberus: (1 

Flyby Mission to Pluto 

AAE 241: Spacecraft Design Proposal 

Professor: K. Sivier 

Section Teaching Assistant: A. Koepke 

University of Illinois 

at Urbana-Champaign 

NASA Advanced Design Program 

Spring Semester 1990 

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https://ntrs.nasa.gov/search.jsp?R=19910006735 2019-05-09T14:35:27+00:00Z
"-,..-

J~-,~ '41 '1l1"'?4 

Theodore W. Humphrey, Project Leader 

Attitude, Articulation, & Control 

~l:!$:~/Y~ i:~ 

Mission Management, 

~ Planning, & Costing 

Bruce E. Hackett 

Science 

2~j ;(I&~ Ral~h G. Kennedy 

Propulsion 

Structures 

2_ 

Command, 

i!!:i~ Control, & Communication 
Section

INTRODUCTION

Section 1-1:

Section 1-2:

Section 1-3:

Section 1-4:

Section 1-5:

References

Appendix

Appendix

S.e.c.tiaa__ 

Section 2-1:

Section 2-2:

Section 2-3:

Section 2-4:

Section 2-5:

Section 2-6:

A-l:

A-2:

References:

Appendix B-I:

Appendix B-2:

Section _;

Section 3-1:

Section 3-2:

Section 3-3:

Section 3-4:

Section 3-5:

References

Appendix C:

Title

TABLE OF CONTENTS

MISSION MANAGEMENT, PLANNING, &

MMPC REQUIREMENTS

SELECTION OF MISSION TYPE

TRAJECTORY DETERMINATION

MMPC EFFECTS ON OTHER SUBSYSTEMS

COSTING

COSTING (MMPC)

ACRONYMS

EQUATIONS AND CONVERSION FACTORS

STRUCTURES

LISTING OF REQUIREMENTS

MAJOR DESIGN FEATURES

LAYOUT OF COMPONENTS/INERTIA PROPERTIES

MATERIAL SELECTION

THERMAL CONSIDERATIONS

LAUNCH VEHICLE COMPATIBILITY AND ON-ORBIT

ASSEMBLY

NUMBERED

OTHER

ITERATIONS OF THE SPACECRAFT CONFIGURATION

MINIMIZATION OF TRUSS MASS

POWER AND PROPULSION

INTRODUCTION

POWER

PROPULSION MODULE

ELV/UPPERSTAGE TRANSFER VEHICLE

CONCLUSION

EQUATIONS

1

3

3

4

11

12

16

16

17

18

19

25

28

31

33

35

35

36

37

39

40

45

51

52

53

54
_ection 4:

Section 4-1:

Section 4-2:

Section 4-3:

Section 4-4:

Section 4-5:

Section 4-6:

References

Appendix D

S.ey..fiaL. 

Section 5-1:

Section 5-2:

Section 5-2.1:

Section 5-2.2:

Section 5-2.3:

Section 5-3:

Section 5-3.1:

Section 5-3.2:

Section 5-3.3:

Section 5-4:

References

Appendix E-1

Section 6:

Section 6-1:

Section 6-2:

Section 6-3:

Section 6-4:

Section 6-5:

Section 6-6:

Section 6-7:

References:

ATTITUDE, ARTICULATION, AND CONTROL (AACS)

INTRODUCTION

MAJOR FEATURES OF THE AACS

HARDWARE SELECTION AND PLACEMENT

SCANNING AND POINTING REQUIREMENTS

IMPLEMENTATION

ATTITUDE CONTROL MODES

CONCLUSION

COMMAND, CONTROL AND COMMUNICATION

INTRODUCTION

COMMAND AND CONTROL

COMPUTER COMMAND SUBSYSTEM

ATTITUDE AND ARTICULATION CONTROL SUBSYSTEM

FLIGHT DATA SUBSYSTEM

COMMUNICATIONS

TELEMETRY

RADIO FREQUENCY SUBSYSTEM

SPECIFICATIONS

CONCLUSION

SCIENCE SUBSYSTEM

RFP REQUIREMENTS

METHOD OF ATTACK

SCIENCE OBJECTIVES

COMPONENTS

SCIENCE TIMELINE

INTERACTION WITH OTHER SUBSYSTEMS

FUTURE CONCERNS

NUMBERED

OTHER

IMPLEMENTATION PLAN

56

57

62

66

66

67

68

69

70

70

72

73

73

73

74

75

76

77

79

81

83

84

85

86

91

93

94

95

95

96

ii
INTRODUCTION

Pluto, the ninth planet in the solar system, is named after the Greek god

of the Underworld. The namesake of this project is Cerberus, Pluto's watchdog

which faithfully stood guard at the gates of Hades.

[ •

Cerberus is designed to meet the requirements stated in the Request for

Proposal (RFP). Those requirements that apply to all subsystems are

summarized below:

1)

2)

3)

4)

5)

6)

7)

8)

9)

Develop a conceptual design for an unmanned spacecraft to study

Plutonian space.

Optimize performance, weight and cost.

Spacecraft should be simple, reliable and easy to operate.

Use off-the-shelf hardware and technology available by 1999.

Identify and minimize on-orbit assembly.

Should be able to perform several possible missions.

Sufficient design lifetime to carry out its mission plus a reasonable

safety margin.

Use latest advances in Artificial Intelligence.

For costing and overall planning, assume that four spacecraft will be

built.

The goal of the Cerberus Project is to design a feasible and cost-effective

mission. The design stresses proven technology that will avoid "show

stoppers," which could halt mission progress. Cerberus also utilizes the latest

advances in the spacecraft industry to meet the stringent demands of a
journey to the edge of the solar system.

means to unlocking the mysteries of Pluto.

The result is Cerberus, a practical
r

Section 1-1:

Section 1: MISSION MANAGEMENT. PLANNING.

& COSTING (MMPC)

MMPC REOUIREMENTS

MMPC entails several requirements from the Request For Proposal (RFP)

specific to the subsystem. These are in addition to general requirements

which pertain to the mission as a whole. The MMPC requirements, specific and

derived, are listed below.

1)

2)

3)

4)

5)

6)

7)

8)

9)

10)

Spacecraft must travel to Plutonian space.

Spacecraft must travel to Plutonian space via an optimal

trajectory.

a.

b.

Trajectory should be optimized for Av.

Trajectory should be optimized for time

of flight (TOF).

Individual bums must remain within limits of

propulsion system.

Spacecraft trajectory should not subject spacecraft to

conditions which will cause it undue damage.

MMPC analyst must perform cost estimates of individual

subsystems.

MMPC analyst must perform cost estimate of entire

mission.

Mission type must be one of three types:

flyby, orbiter, or lander.

MMPC analyst must outline mission sequence of events.

Launch should take place sometime between 2000 2010 A.D.

Acronyms for Section 1; MMPC are listed in Appendix A-1.

Section 1-2: SELECTION OF MISSION TYPE

The flyby type of mission has been determined to be the best for the

Cerberus project. This decision was based upon the results of preliminary

trajectory studies using the multiple impulse optimizing program, MULIMP, as

well as other considerations. These considerations include:
1)

2)

3)

4)

Section 1-3:

Existing technology does not facilitate the

transportation of fuel mass necessary to burn into

orbit capture at Pluto. This precludes the orbiter and

lander class missions. MULIMP studies produced Av values which

would be required for orbit capture. There exists no trajectory

which would allow orbit capture at Pluto, given the available

technology, and the acceptable TOF for this mission.

Pluto and Charon are only separated by 19400 km.[3] An object

placed in orbit about Pluto would likely have its orbit perturbed

such that it left Plutonian space.

Other general solar system science can be emphasized

on a flyby mission. Examples are a Jupiter study

before arrival at Plutonian space, and a measurement

of the heliopause after Pluto passage.

This is a preliminary mission; it is prudent for this to

be of the flyby class given present knowledge of

Plutonian space. This mission will assist in

determining the benefit of another Pluto mission. This

venture will yield information required for launch of

an orbiter or lander Class mission, when technology for

such a mission becomes available. It is also prudent to

keep this mission at the flyby class to keep from

overshadowing the Lunar and Mars initiatives.

TRA.IECTORY DETERMINATION

Several paths were studied as a possible route to Pluto.

included different gravity assist flyby trajectories:

1)

2)

3)

4)

5)

Earth Pluto

Earth Jupiter- Pluto

Earth - Mars - Jupiter- Pluto

Earth - Venus - Earth - Earth - Jupiter - Pluto

Earth - Earth - Jupiter - Pluto

These paths

4
i

Certain data were required to facilitate logical study of trajectories with

MULIMP. Values of ,_v and TOF required for Hohmann transfer between the

various planets of interest were calculated by hand. These are listed in Table

1-1. Equations and parameters used in these calculations are listed in

Appendix A-2.

Table 1-1: Hohmann Transfer Values

Planets of Interest

i

Earth & Pluto

Earth & Mars

Earth & Jupiter

Av_km/s}

11.8120

2.9458

8.7920

TOF(days)

16581.5465

258.9324

997.5984

Mars & Jupiter 5.8968 1125.6354

The orbit periods and synodic periods of the planets of interest are

given in Table 1-2.[2]

Table 1-2: Orbit Periods and Synodic Periods

Planet

Venus

Mars

Jupiter

Pluto

Orbit Period

224 days

687 days

11.9 years

247 years

Synodic

Period_da],s_

584

778

398

367

The first path to be investigated was a direct transfer from Earth to

Pluto. Observation of hand calculated data and study of available literature[3]

lead to the following conclusions:

1) The mission TOF would be too long for practical

purposes using existing technology.
2)

3)

The zXv at Earth parking orbit (1.0437 Earth radii)

would be too large for available propulsion systems.

The launch energy (C3) would be in excess of the

capabilities of existing launch vehicles.

The second path to be investigated was a transfer to Pluto with a gravity

assist at Jupiter. The first step in studying this trajectory was to use MULIMP

to find the optimal transfer from Earth to Jupiter, in terms of zXv, in the early

part of the first decade of the twenty-first century. Knowing this locally

optimal launch date, MULIMP optimized the trajectory to its completion at

Pluto. The first launch date was then incremented by an amount equal to

Jupiter's synodic period, and the Earth - Jupiter - Pluto transfer was again

optimized by MULIMP, This process was repeated until the incremented

launch date fell outside the ten year launch window prescribed by the RFP.

MULIMP data also displayed position of each event in three dimensional

Cartesian coordinates, with the sun as origin and its ecliptic as the X-Y plane.

With this data, a path could be graphically plotted to ensure smooth flow of the

trajectory in a counter-clockwise manner.

Upon examining the various data output by MULIMP, it was clear that

although the Earth Jupiter - Pluto trajectory was an improvement upon the

Earth Pluto trajectory, it was still unsatisfactory for these reasons:

1)

2)

3)

The TOF was generally greater than 18 years.

The z_v at Earth parking orbit was larger than desired

for cost effectiveness of the mission.

The launch energy approached or exceeded

unacceptable levels (i.e., from 90 to several thousand).

The next path under consideration was Earth - Mars Jupiter Pluto. This

trajectory was studied employing a similar method of attack to that used for the

Earth Jupiter Pluto study. The best early launch date to Mars was

determined using MULIMP, and this launch date was incremented by Mars'

synodic period through 2010. At each launch date, MULIMP optimized the

complete Earth - Mars - Jupiter - Pluto transfer. Several MULIMP studies were

also performed in the region of time of two months preceding and following
each of the aforementionedlaunch dates. Once again, this trajectory displayed

improvement over the previous path studied, particularly in these areas:

1)

2)

3)

Launch energy: C3 fell generally in range of 10 - 15.

Av from Earth parking orbit was reduced to the range

of 3 to 4 krn/s

TOF was in the range of 15_+2 years.

i

Prohibiting problems with this flyby path arose at Mars, where an

impulse of at least 6.3 km/s was required.

The ensuing path under scrutiny was an extension of Galileo's Venus

Earth - Earth Gravity Assist (VEEGA) trajectory.J4] A flyby of Jupiter was added

before the spacecraft continued on to Pluto. It became apparent upon

reviewing MULIMP test output that the planets of the solar system are not in

position conducive to this type of transfer during the prescribed ten year

launch window. Although TOF was reduced significantly to roughly six years,

flyby impulses unattainable.

The final trajectory to be considered was an Earth- Earth Jupiter -

Pluto path. The method of attack for studying this path was again similar to

that described previously. This trajectory yielded the most satisfactory values.

Table 1-3 displays values from each type of trajectory studied. These values

were compared in a trade study manner to determine the best course of flight.

Although these values are not likely to be the absolute optimum value for each

case, they were determined to be sufficiently representative.

Table 1-3: Characteristic MULIMP Values for Different Trajectories

i

EP

Path

EJP(typical)

EJP(unusual)

Av(km/s)

N/A

11.195

8.194

TOF(_ears)

N/A

26.935

15.383

C3

N/A

94.560

90.523

Comments

Excluded from

MULIMP study

Path falls through

Jupiter at .8 planet

radii
EMJP(typ.)

EMJP(unus.)

EVEEJP

EEJP

E: Earth

10.915

3.788

93.556

5.941

13.665

14.727

6.072

18.688

J: Jupiter M: Mars P: Pluto

11.376

10.814

13.687

47.518

V: Venus

AV at Mars: 7.201

km/s.

Unattainable.

Path falls through

center of Mars & .723

Jupiter radii.

Combined Av for the

two E flyby's: 89.733

km/s.

The Earth

considerations:

Earth - Jupiter - Pluto trajectory was selected due to several

1)

2)

3)

4)

5)

6)

The Av from Earth parking orbit was determined to be

attainable for Cerberus' mass (see Section 3:

Propulsion).

The launch energy was determined to be attainable for

this mission using existing technology (see Section 3:

Propulsion).

Midcourse and flyby hv's were determined to be

feasible(see Section 3: Propulsion).

Flyby of Jupiter occurs at a distance which does not

require addition of radiation shielding to Cerberus'

structure.

There is a launch window of eleven days.

Launch occurs early enough in the 2000 - 2010 period

to allow postponement and still make that ten year

window.

A later launch date would allow for more development time, better

weather for launch, and a chance to find a better trajectory. However, Pluto is

moving away from its perihelion distance of 29.6 AU, which it reached in 1989.
Therefore, an earlier launch has the potential to travel less distance to

Plutonian space.

Table 1-4 displays information on Cerberus' trajectory, and outlines the

mission sequenceof events. The launch date shown falls in the middle of the

eleven day launch window.

Date

2002 Jan4

Table 1-4: Cerberus' Traiect)rl¢ & Sequence of Events

Av Radius of Passage

Ikm/st

5.195

2003 Jun 11 0.418

2004 Nov 25

2005 Jan 23

2006 May27

2006 Jun 21

2020 Aug 18

Event

Depart Earth parking

orbit

Midcourse burn

Earth flyby

Midcourse burn, Plane

change

Begin Jovian science (50

days)

Jupiter flyby:

closest approach

Begin Plutonian science

(50 days)

Pluto flyby:

closest approach

2020 Sep 12

0.008

0.320

0.000

1.0437 planet radii

1.3000 planet radii

26.780 planet radii

The redeeming values of this course were considered advantageous

enough to outweigh the disadvantage of the long TOF. The long time is a

disadvantage since the spacecraft, its instruments, and components have not

been tested and proven for such a length of time.

Figure 1-1 graphically displays A v versus launch date in 30 day

intervals for a span of 420 days surrounding the launch date. Figure 1.2

graphically displays A v versus launch date in one day increments for a span

of two weeks surrounding the launch date.

9
60

5O

43

E 33

> 23

1 I I I

0 103 200 300 403

Launch date + 2452000 (Julian Days)

Figure 1-1" Comparison of Delta v versus

Launch Date over 420 days

!

5(30

12

11

E 10

> 9"

6

" I " I ' I " I " I " | " I " I " I " I

70 72 74 76 78 80 82 84 86 88 gO

Launch Date + 2452200 (Julian Days)

Figure 1-2: Comparison of Delta v versus

Launch Date over 2 weeks

10
Section 1-4: MMPC EFFECTS ON OTHER SUBSYSTEMS

ri

MMPC decisions naturally affect the other subsystems involved in the

Cerberus mission, as all of these subsystems are integrated to accomplish one

task.

The science analyst's decisions are influenced by the trajectory.

Cerberus' course of flight determines which objects in space are available for

study. Decisions are further affected by the specific proximity of the

spacecraft to the aforementioned objects during flight, as experiments can be

affected by closeness or distance to the object of scrutiny. The science analyst

must determine from the trajectory the window of time available for study. In

Cerberus' case, there are fifty days allotted for study at Pluto.

TOF is a concern of all functional subsystems: science, propulsion,

attitude articulation & control, communication, and structures. If TOF exceeds

the known service lifetime of a given unit, there may be concern for the

ability of the unit to accomplish its ultimate task.

Closeness of passage to radiating bodies such as our sun and Jupiter is a

structural concern. If flyby is too close to a radiating body, extra shielding

must be added to the craft to protect it from undue damage. In Cerberus' case,

the trajectory does not carry it close enough to the sun to merit unusual

concern. Cerberus' path also falls far beyond the 'very safe' distance of ten

planet radii when flying by Jupiter. The structure must also support the fuel

mass determined by the _v required.

The communication analyst must know when the craft will be in

occultation behind a body. This knowledge is required in order to prepare

autonomous control during this period without contact with the spacecraft.

The propulsion analyst must make decisions for that subsystem based

upon data furnished by MMPC. Launch vehicle and propulsion system

selection must reflect the needs stated for the trajectory. The spacecraft must

carry with it the capability to perform midcourse burns when necessary.

11
Section 1-5: COSTING

i

The estimated cost of the Cerberus mission is $1,069,152,990.00 in

February 1990 dollars. Costing data are displayed in Table 1-6 (Costing

Spreadsheet). Costs are broken down under three headings:

1) Development Project Flight Hardware

2) Development Project - Support Functions

3 ) Flight Project

The costs are evaluated for each unit and are totalled to produce the

mission labor cost and the mission total cost. The sources of the equations and

conversion factors used to determine these values are listed on the Reference

page.[6,7] The conversion factor from FY77 producer dollars to February 1990

producer dollars is 1.815 (see Appendix A-2).[1,5] Mission costs by category are

displayed as a bar graph in Figure 1-3.

CERBERUS MISSION COST BY CATEGORY

DATA AN.

FLIGHT OPS

PGM MGMT

SCI DATA DEV

L + 30 OPS

SYS SUPP

RSI

PART & FLD

LSI

RTGPOV__R

COM & DATA

ANTENNAS

TELECOM

AACS

PROPULSION

THERML CTL

STRUCTURES

I

/

/

[]

I

I I I I I

0 5(_0 1(33333 1,5(3E_ 202£_ 2513300

Figure 1-3: Category Cost - 1990 ($1000)

12
r

Number of Spacecraft: 4

DEVELOPMENT PROJECT - FLIGHT HARDWARE

FLIGHT HARDWARE TOTAL COST: $572892850.01

STRUCTURES AND DEVICES

MASS (kg) 168.00 ADJUSTED

DLH 419.99 262.49

RLH 178.15 178.15

SYSTEM SUBTOTAL: 440.64

THERMAL CONTROL, CABLING, AND PYROTECHNICS

MASS (kg) 82.00 ADJUSTED

DLH 276.86 69.22

RLH 126.47 126.47

SYSTEM SUBTOTAL: 195.68

PROPULSION

MASS (kg) 90.00 ADJUSTED

DLH 596.57 551.82

RLH 201.13 201.13

SYSTEM SUBTOTAL: 752.96

ATTITUDE ARTICULATION AND CONTROL

MASS (kg) 85.00 ADJUSTED

DLH 1202.55 1112.36

RLH 656.88 656.88

SYSTEM SUBTOTAL: 1769.24

TELECOMMUNICATION

MASS (kg) 10.00 ADJUSTED

DLH 189.44 189.44

RLH 130.37 130.37

SYSTEM SUBTOTAL: 319.80

ANTENNAS

MASS (kg) 40.00 ADJUSTED

DLH 1147.07 1147.07

RLH 534.24 534.24

SYSTEM SUBTOTAL: 1681.31

COMMAND AND DATA HANDLING

MASS (kg) 15.00 ADJUSTED

DLH 186.03 186.03

RLH 90.33 90.33

SYSTEM SUBTOTAL: 276.37

Table 1-6: Cerberus Mission Costing Spreadsheet Values (page 1 of 3)

Note: Labor Hours listed in 1000's of hours

Dollar amounts given in February 1990 producer dollars

13
l

MITG POWER

MASS (kg) 50.00 ADJUSTED

DLH 547.61 342.26

RLH 348.15 348.15

SYSTEM SUBTOTAL: 690.40

LINE SCAN IMAGING

MASS (kg) 29.70 ADJUSTED

DLH 761.60 761.60

RLH 450.03 450.03

SYSTEM SUBTOTAL: 1211.64

PARTICLE AND FIELD INSTRUMENTS

MASS (kg) 28.00 ADJUSTED

DLH 692.84 692.84

RLH 577.58 577.58

SYSTEM SUBTOTAL: 1270.42

REMOTE SENSING INSTRUMENTS

MASS (kg) 27.00 ADJUSTED

DLH 305.49 305.49

RLH 40.20 40.20

SYSTEM SUBTOTAL: 345.69

TOTAL HARDWARE ADJUSTED DLH

TOTAL HARDWARE ADJUSTED DLH + RLH

5620.61

8954.15

DEVELOPMENT PROJECT - SUPPORT FUNCTIONS

SUPPORT FUNCTIONS TOTAL COST: $180416000.00

SYSTEM SUPPORT AND GROUND EQUIP.

DLH 1732.94

LAUNCH + 30 DAYS OPERATIONS AND GROUND SOFTWARE

DLH 551.27

SCIENCE DATA DEVELOPMENT

DLH 108.25

PROGRAM MANANGEMENT/MA&E

DLH 601.40

Table 1-6: Cerberus Mission Costing Spreadsheet Values (page 2 of 3)

Note: Labor Hours listed in 1000's of hours

Dollar amounts given in February 1990 producer dollars

14
l

I

FLIGHT PROJECT

FLIGHT PROJECT TOTAL COST:

FLIGHT OPERATIONS

DLH 3544.50

DATA ANALYSIS

DLH 1506.41

$180416000.00

TOTAL MISSION DLH

TOTAL MISSION DLH + RLH

MISSION TOTAL COST:

13665.39

16998.93

$1069152990.01

Table 1-6: Cerberus Mission Costing Spreadsheet Values (page 3 of 3)

Note: Labor Hours listed in 1000's of hours

Dollar amounts given in February 1990 producer dollars

15
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i

.

.

3.

o

°

.

.

References

Carson, C. S. (editor in chief), Survey Qf C_rren_ Business, Volume 70,

Number 2, February 1990

Conway, B. & Prussing, J., Orbital Mechanics, 1989, p. 127

Croswell, K., Pluto: Enigma on the Edge of the Solar System,

Astronomy, July, 1986

D'Amario, et al., Galileo, 1989 VEEGA Mission Description: Paper AAS

89-431, JPL AA$/AIAA Astrodynamics Specialist C_)nfcren_e, Stowe,

VT, August 7-10, 1989

Kincannon, C. L., et al., Statistical Abstract of the United States,109th

Edition, United States Bureau of the Census, Washington, D. C., 1989

Kitchen, L. D., Manpower Cost Estimation Model for Automated

Planetary_ Projects - 2, Report No. SAI 1-120-399-C2, Science

Applications, Inc., September, 1973

Pekar, P. et al., Manpower Cost Estimation Model for Automated

Planetary Projects, Science Applications, Inc., September, 1973

Appendix A-l; Acronyms

AU

Av

DLH

km

L

MMPC

MULIMP

RFP

S

TOF

VEEGA

Astronomical Unit (149.6E6 km)

change in velocity (km/s)

Direct Labor Hours

kilometer

Launch

Mission Management, Planning, and Costing

Multiple Impulse Optimizing Program

Request for Proposal

second

Time of Flight

Venus - Earth - Earth Gravity Assist

16
Aanendix A-2: Eouations and Conversion Factors

Hohmann Transfer

rVenus = .723 AU

rEarth = 1 AU

rMars = 1.52 AU

rJupiter = 5.203 AU

rpluto(average) = 39.4 AU

rpluto(perihelion) = 29.6 AU

IXSun = 1.327Ell km3/s 2

R=r2/rl

AVl/Vcl = (2R/(I+R)) "5 - 1

VCl = (_Sun/rl) "5

Av2/Vcl = R (-'5) (2/(R(I+R))) "5

a = (rl + r2)/2

tHohmann = _(a3/p.Sun) •5

Costing

Purchasing Power of Dollar

FY77 1982/1984 February 1990

Consumer 1.649

Producer 1.546

Consumer Dollar Conversion Factor:

Producer Dollar Conversion Factor:

1.00

1.00

2.1088

1.8150

.782

.852
Section 2: STRUCTURE_

Section 2-1: LISTING OF REOUIREMENTS

Easily the most difficult part of designing any spacecraft is dealing with

the vague, and often contradictory, requirements of the mission.

Requirements such as the ones listed in the Request for Proposal (RFP)

identify the objectives of the design, and it is up to the analyst to achieve the

optimum solution. Listed at the beginning of this proposal are the

requirements that must be met by all subsystems. Concepts such as

minimizing cost, keeping the design simple and reliable, etc., must be on the

mind of the analyst at all times. The most important part of preliminary

design is meeting as many, if not all, of the requirements outlined by the RFP.

As well as meeting the overall objectives of the mission, each subsystem

must also satisfy many derived requirements. These derived requirements are

based on the objectives outlined in the RFP, but they are specific to the

subsystem.

For the Structures Subsystem, the derived requirements have a great

deal to do with the overall design of the spacecraft (spacecraft). Below is a

listing of the derived requirements for the Structures Subsystem. All of them

are objectives specific to this subsystem, but they are based on concepts

outlined in the RFP.

1)

2)

3)

4)

5)

6)

7)

8)

9)

Overall design of the Cerberus spacecraft

Design to maximize science performance

Calculation of inertia properties of spacecraft

Material selection for the various components

Thermal control considerations

Verification of Launch Vehicle Compatibility

Identification of any On-Orbit Assembly (OnOA)

Structural analysis of truss bays to meet launch conditions

Identification of subsystem interaction

The RFP requirements kept in mind when designing the spacecraft as

well as the overall mission plan agreed on by the group. In essence, Cerberus

is meant to be a conservative project. With the Moon/Mars initiative the

18
f :

i

centerpiece of both NASA's and Congress' space commitment, it was believed

that this was the best approach. Nothing was to be done that would

overshadow these important advances. From the beginning, Cerberus was

meant to be an inexpensive mission. This meant that as much off-the-shelf

hardware should be used. Fortunately, with the Mariner Mark II (MMII)

program just beginning, it was believed that a good amount of off-the-shelf

hardware would be available.

This conservative, off-the-shelf approach is well reflected by the

structural design methodology. Older missions such as Voyager and Galileo

were studied to understand what problems they had and the solutions to those

problems. The more recent MMII program gave valuable insight into new

methods of spacecraft design. The most important part of studying these

previous missions was understanding how the mission to Pluto differed. With a

mission lifetime of 18.7 years, and a safety margin of approximately 5 years, it

is easy to see the major differences between the projects. Only Voyager has

come close to having a mission lifetime of this magnitude. The conservative

basis of Cerberus is well-suited to meeting this difficult requirement. Material

selection, spacecraft configuration, and thermal design all reflect this overall

mission plan.

The most important requirement for the structural designer was

ensuring Cerberus ability to carry out its mission. The objective of a Pluto

probe is to gather as much scientific data as possible, in the most efficient and

cost-effective way. This must be kept in mind at all times when designing the

spacecraft. It is believed that this design, with its conservative and off-theshelf approach, is the most effective and cost-efficient way to successfully

complete the long journey to Pluto.

Section 2-2: MA.IOR DESIGN FEATURES

The most important part of the Structures subsystem is determining the

overall layout of Cerberus. There is a great deal of system interaction that

occurs for an effective design. Not only must the requirements of the RFP be

satisfied, but any constraints imposed by the different subsystems must be

taken into consideration.

Figure 2-1 represents a view of Cerberus from the bottom of the

spacecraft (see next page). Some of the major design features are visible from

19
(_CM of spacecraft

4.8m meter

Hydrazine

Bladder (I of 2)

Th!uster

(I of 4)

MITG (2)

i0 Bay Bus

(i.5m wide)

Scale: 1/25

4cm = I m

Figure 2-I Bottom view of

Cerberus

ZU

High Accuracy

Science Platform

Antenna

Heliu

Tank

(i of 2)

Mag

Cannister

When dep

om wil

I

Ted,

_xtend

Low Accuracy

Science Platform
r'-

,f :

k

this angle. A 10 bay bus is the major structural component of the spacecraft.

It contains the electronics that control the mission to Pluto. The following is a

list of the major design features, grouped by subsystem.

Table 1-1"

Science

CCC

Summary of major design features

Cerberus contains two science platforms, 1 for

High Accuracy science (HAP), the other for

Low Accuracy science (LAP). Both platforms

have a good field of view, maximizing the

science performance of the instruments.

The HAP is protected by a radiation shield. This

will minimize the effects of the space

environment by protecting the instruments

from both radiation and micrometeroid impact.

The MITGs have been placed opposite the

magnetometer to eliminate interference.

14.5 m separates the two components.

Over

A dipole antenna has been placed on Cerberus

for radio science. The two antennas were

placed at a ,90 angle to each other to satisfy

science requirements.

A 4.8 m High Gain Antenna (HGA) is the

component that is responsible for telemetry. It

has been placed on top of the 10 bay bus. It can

be folded for launch.

21
Power &

Propulsion

Attitude,

Articulation and

Control

Mission

Management and

Planning

Two MITGs are the main power source for

Cerberus. They have been placed on a boom .8

m away from the bus.

A propulsion module that will provide

necessary Av's for the spacecraft is attached at

the bottom of the bus. It's consists of 2 bladders,

2 tanks and a support structure. Two large

bladders (r=.335 m) are for the Hydrazine

propellant, two small aluminum tanks (r=.ll0

m) contain the Helium pressurant. A plate

made of titanium is the material for the support

structure.

There are 4 sets of thrusters that control the

attitude of the Cerberus spacecraft. They are

placed along the principal axes of the craft•

• A star tracker and sun sensor have been placed

on the HAP for inertial reference.

• The dimensions of the spacecraft have been

sized to conform to the launch vehicle•

Since the scientific instruments will be fully operational during the

whole flight, there is nothing in the design that precludes it from performing

several possible missions.

The overall design approach reflects the conservative nature of the

mission. Most of the equipment is off-the-shelf hardware, with inheritance

from both Galileo and the more recent MMII program• Cerberus is simpler

than Galileo, with no spun sections, and only one bus. Most of the structural

design reflects the newer MMII program [1]. The 10 bay bus was chosen to

take advantage of any extras from this program, since the Comet Rendezvous

and Asteroid Flyby (CRAF) uses this configuration. The two scan platforms also

22
reflect the CRAF inheritance, as does the short MITG boom and the long (12m)

magnetometer truss structure. All of these design features represent

compliance with the requirement to maximize use of off-the-shelf hardware.

Figure 2-2 representsa side view of the Cerberusspacecraft. This

angle shows two features that differ from the CRAF configuration. The first is

the Galileo type antenna (4.8m diameter) that will provide telemetry for the

mission. Again, even this feature is inherited from the Galileo project, so it

does not represent a major redesign. The propulsion module also represents a

change from MMII. It is smaller, fitting inside the 1.5m space inside the bus.

Although different from MMII, it is not much different than modules used for

other interplanetary missions.

As mentioned before, the 10 bay bus contains the electronics that will

control the mission to Pluto. Each bay contains electronics for one subsystem,

although each subsystem was assigned more than one bay. Figure 2-3 is a

graphical representation of each of the bay assignments.

7 9 ")

Figure 2-3: Electronics contained in each bay

The bay assignmentswere selected to minimize cabling and also to even

out the inertia properties of the electronics stored in the bus.

23
Section 2-3: LAYOUT OF COMPONENTS/INERTIA PROPERTIES

Now that the major design features of the Cerberus have been discussed,

the exact layout of the components and the inertia properties will be studied.

The location of the components is directly related to the inertia

properties of the spacecraft. In turn, the problem of attitude control is

heavily dependent on the inertia properties. To make the attitude control

problem as simple as possible, there were certain objectives of the design.

These objectives are listed below.

1) The center of mass of the whole structure should be kept near the

middle of the bus. A good location was selected at (0,0,-.40).

2) The off-diagonal terms of the inertia tensor should be kept at a

minimum. This means that the principal directions of the

spacecraft lie along the directions chosen for the initial layout (See

Figures 2-1 and 2-2 for these axes).

An initial configuration for Cerberus was drawn up. In the

preliminary phase, close attention was paid to the symmetry of the spacecraft.

The two science platforms were placed on opposite ends of the bus since it was

believed that their inertia properties would even out. The magnetometer and

the RTGs were then placed at 90" angles to these platforms. Constraints

imposed on the design were kept in mind at all times. Most importantly, the

RTGs were kept far from the magnetometer, and the HAP and the LAP were

given a good field of view. The size of the propulsion module enabled it to be

placed inside the 1.5m wide electronics bus.

Two iterations resulted in the final placement of the spacecraft

components. The inertia matrices for each of these iterations, as well as the

center of mass (CM) and principle directions, are given in Appendix B-1. Only

the final inertia tensor will be given here. A summary of how the

configuration changed is below in Table 2-2.

25

pRECEDING PAGE BLANK NOT FILMED
Table 2-2:

Initial

Design

After 1st

iteration

After 2nd

iteration

Summary of Configuration Changes

The off-diagonal terms of the inertia

matrix were small and the

eigenvectors were close to the chosen

axes. The CM was 25 cm too far in the

x-direction and 12 cm in the ydirection.

The LAP was moved .5m further away

from the spacecraft and the HAP was

moved .6m closer to the bus. The

inertia matrix and eigenvectors were

still good. The CM was still l lcm too

far in the x-direction and 9cm too far

in the y-direction.

The MITGs were moved .5m closer to

the bus. 5 kg of mass (considered

thermal protection) was added to the

LAP. This resulted in the final inertia

properties.

The propellant weight was added to the bladders and the total inertia of

Cerberus plus Hydrazine was obtained. The principle inertia matrix is given

below.

Principle

Matrix

Inertia

2192 0 0 0 886 0

0 0 2800

J

I

The location of the CM is shown in Figures 2-1 and 2-2. Once the CM and the

eigenvectors were found, the thrusters were placed. The thruster placement is

also shown the above mentioned figures.

The final inertia properties of Cerberus represent a satisfactory

configuration. Both the principal directions and the CM are well-placed,

26
allowing for easy attitude control. Also, the principal inertias are not outside

the range of normal thruster sizes. The final configuration of the spacecraft

allows for precise attitude control, thereby increasing the science

performance of the flyby mission.

One configuration problem that needed to be addressed was "center of

mass migration" (author's term). This problem is the result of propellant loss

during the mission. As the Hydrazine bladders empty, the CM will "migrate"

away from its original position. This could cause an attitude and control

problem during the flight or, more importantly, at Pluto rendezvous. This

effect was studied by calculating the inertia properties of Cerberus at varying

stages of propellant loss. The beginning of the mission was assumedto have

0% loss, the end 100%. It was found that the principal inertias changed by

2.4%, which is not a significant amount. A more important change occurred

in the location of the CM. The results are plotted below in Figure 2-4.

Change in CM due to Propellant Loss

of

-.0.2

0 93 4O 6O 80 100

Percent Loss

Delta Xcm

Delta Ycm

Delta Zcm

Figure 2-4: Center of Mass "Migration"

Figure 2-4 shows that the parameter most affected by propellant loss is the z

component of the center of mass. The negative sign in AZcm appears because

i

27
!

T

of the chosen direction for the z-axis. The CM is moving 'up' the spacecraft, by

as much as 17.53cm for empty bladders. This could cause attitude control

problems as the fuel is expended.

A simple solution to this problem exists. Since all four of the booms

have the ability to move up and down, due to the launch requirements (see

Section 2-6), they can be used to counteract this mass loss. By moving both the

MITG and magnetometer booms .5m in the positive z direction, the CM of mass

can be placed only 6cm away from its original position. This also assumes that

30% of the propellant is left in the bladders, which is slightly over the excess

allotted for the mission. It may be possible to do even better by lowering the

science platforms, but this is not desirable. Moving the science platforms

might cause a loss in pointing accuracy, which in turn would degrade science

performance. A 6cm movement in the CM is much better performance than a

17.5cm shift. This is not expected to cause any attitude and control problems.

Section 2-4: MATERIAL SELECTION

One of the most important requirements of spacecraft design is weight

minimization. Weight minimization is dependent on the materials selected for

the spacecraft. Ideally, a designer will choose the lightest available materials.

Unfortunately, this is not the only factor involved. The materials must be

space-proven and reliable, thereby maximizing the chances for a successful

mission. Depending on the application, the material must have a high yield

strength, so it will not fail at launch. All these requirements must be taken

into account when selecting the materials for the mission to Pluto.

The first stage to selecting materials was to gather important properties.

Table 2-3 is a list of commonly used space materials and their relevant

properties. [2]

28
D- k

Table 2-3: Properties of materials for space use

Material

A1-2024

A1-2219

Beryllium

Mg-Hm21A

Stainless Steel

Ti-6AI-4V

Graphite Epoxy

Graphite A1

density

_kg/m 3)

2801.55

2829.29

1858.45

1775.24

7933.10

4549.05

1941.67

3051.19

yield

stren gth

(Mea)

375.20

369.33

381.06

193.46

1084.55

785.57

1055.24

586.25

_P___

_y

7.47

7.66

4.88

9.18

7.31

5.79!

1.84

5.20

Table 2-3 displays the varying material properties. These properties are

not the only factor that must be taken into account when selecting materials.

Given the conservative, low cost nature of the Cerberus mission, it is very

important to look at the reliability factor. The most commonly used material

for space applications is aluminum. It is easy to form and has proven its

worth.J2] Titanium is also space-proven, as well as steel. Many of the other

materials in Table 2-3 do not have the reliability and ease of fabrication that

the above materials provide. The composites are very expensive and are

subject to degradation in UV environment.J2] Berylliums are difficult to form,

and are toxic. Magnesiums are difficult to weld, which increases fabrication

COSt.

In keeping with the overall Cerberus objective of designing a

conservative, inexpensive spacecraft, the materials selected are all spaceproven and relatively easy to fabricate. Table 2-4 summarizes these decisions.

29
Table 2-1" Material Comparison

¢

Aluminum

Titanium

Steel

The material that has been the mainstay of

spacecraft. Will be used for the 10 bay bus and

most of the truss elements. Also used for science

platforms and the radiation and MITG shields.

Given its lower weight to strength ratio, titanium

will be used wherever load is carried. This

includes the support plate for the propulsion

module and some of the truss members.

Although heavier than aluminum and without the

excellent weight to strength ratio of titanium, steel

will still be used for pins, springs, etc.

bay bus.

The most important trade-off involves selecting a material for the I0

Figure 2-5 displays the relative masses of a 1.5m wide,10 bay bus.

Material selection for 10 bay bus

z103-

_ bus

62.5 % lighter

_ than Titanium

0

Aluminum Titanium Steel

Figure 2-5: Material Selection for 10 bay bus

The aluminum bus definitely satisfies the minimum weight

requirement, being 62.5% lighter than a titanium bus of the same

configuration. Steel is not even a consideration due to its weight.

Selection of aluminum is also in keeping with the requirement for

simplicity and reliability. It is both space proven and easy to fabricate. It can

30
easily be considered an off-the-shelf item, which again decreasescost of the

overall mission.

In material selection for the truss members, the decision was made to

use both aluminum and titanium. Aluminum will be used in the members that

do not have to carry the high launch loads, while titanium will be the material

for the truss elements that do. For a complete discussion of this analysis, see

Appendix B-2.

Section 2-5: THERMAL CONSIDERATIONS

Given the length of Cerberus' mission to Pluto, the thermal control

problem is one that needs to be studied closely. In general, there are two types

of control that could be used: passive and active. Passive involves the use of

insulating blankets and louvers to reduce the escape of heat from a

component. Active control comes in two forms, electric heaters and

Radioisotope Heating Units (RHUs). Table 2-4 summarizes the disadvantages

and advantages of these three types of thermal control.

Table 2-4: Design Trade-Offs for Thermal Control

Type

Louvers

Mylar

Insulation

Electric

Heaters

RHU

Advantages

Used to emit heat from

electrical components.

Insulation used to reduce

heat loss from a

component. Also good for

micrometeroid protection.

Can control temperature

over relatively large

range.

Uses no electrical power.

Disadvantages

Only good for emitting

heat.

Passive control system,

might not last the complete

mission.

Uses electrical power.

Supplies only 1 Watt of

power. Need to be used in

quantities.

31
The thermal control of Cerberus will involve a combination of all the different

types listed in Table 2-4. The problem has been broken down into three areas:

the electrical bus, the science platforms, and the propulsion module. Table 2-5

lists the thermal control for each area.

10 bay bus

LAP and

HAP

Propulsion

Module

Table 2-5: Thermal Control for Cerberus

A combination of heaters and Mylar insulation will be

used. Louvers will be placed on the boom side of the bus

to enable heat emissivity from the electrical components.

For bays with no heat storage, Mylar insulation will be

used to reduce heat loss.

Three types of control are necessary. When the

instruments are not operating, electrical heaters will be

used for temperature control. Mylar insulation will be

used to reduce heat loss and for micrometeroid protection.

Finally, louvers are placed on the 'inside' of the platform

to allow heat to escape during periods of high instrument

use.

Mylar blankets can be placed over the bladders and tanks

to reduce heat loss. RHUs will be used (approx. 50 of

them) for thermal control during flight. These will be

placed on the support structure of the module.

't .:

This thermal control design is again reflective of the overall nature of the

mission. Inheritance from MMII program can be seen [3], thereby increasing

the use of off-the-shelf hardware. The system is redundant, especially in the

important area of the science platforms. This is needed, given the length of

the mission to Pluto.

One design change was looked into. This involved the use of waste heat

from the RTGs for thermal control of the propulsion module. Two factors

excluded this design. The first was that the propulsion module was small

enough to fit inside the bus, thereby making it more efficient to use RHUs.

The second problem was the fact that the RTGs will be moving down during the

mission for inertia control. This was discussed in Section 2-3.

32
Section 2-6: LAUNCH VEHICLE COMPATIBILITY AND

ON-ORBIT ASSEMBLY

(--

To ensure that the Cerberus spacecraft would be compatible with the

launch vehicle, three measures had to be taken. First, the bus could not be

overly large. Second, the antenna must be foldable, like Galileo's. Finally, the

booms must be movable to allow them to be folded down.

Figure 2-6 (see next page) shows Cerberus ready for launch. In this

configuration, the spacecraft measures slightly under 3.4m from side to side

and 4.4m from top to bottom. These dimensions are acceptable for modern

launchers. The magnetometer boom has been retracted and the trusses have

been folded down. These will be extended after the spacecraft begins its

journey to Pluto. The truss that mates Cerberus with the launch vehicle will

use explosive bolts to be jettisoned from the spacecraft after disengaging with

the upper stage. It will not be carried along to Pluto.

It is obvious from the diagram that no on-orbit assembly (OnOA) is

required. This is a major simplification to the overall mission. Although the

deployment of the Space Station Freedom in the coming decade makes OnOA a

possibility, there is no need for it. Adding the burden of OnOA only makes the

mission more expensive and complicated. Making Cerberus compatible with

existing launch vehicles and eliminating the need for OnOA makes the design

reliable, more simple, and less expensive.

33
Antenna is folded

and booms are

retracted for

launch

\

)=======

Figure 2-6: Cerberus in launch configuration

34
.

.

.

.

.

Numbered References

Draper, Ronald, "Comet Rendezvous Asteroid Flyby First

MarinerMark II", Advances in Astronautical Sciences. v64 Part I:

399-418.

PP

Koepke, Andy, "AAE 241 Lecture Notes from 2/22/90", Table of Material

Properties

Tsuyuki, G. et al, "Thermal Design for the Comet Rendezvous

Asteroid Flyby Spacecraft", AIAA Paper 89-1753, from AIAA 24th

Thermophysics Conference.

"JPL, Division 35 Mass Estimates", Interoffice Memorandum, Nov. 24

1980, Jet Propulsion Laboratory.

McGill, David, and King, Wilton, Engineering Mechanics:

Introduction to Statics and Dynamics, PWS Engineering

(Boston), 1985.

an

Other References

Galileg: Exploration of Jupiter's System, NASA SP-479

Avila, A, et al, "Thermal Design of the Galileo Bus and Retro

Propulsion Module", AIAA Paper 89-1749, AIAA 24th

Thermophysics Conference.

35
A nnendix ]_-I:

ITERATIONS OF THE SPACECRAFT CONFIGURATION

,4 ,

As mentioned in Section 2-3, there were two iterations done on the

initial configuration. The data for those calculations will be given here. A

complete summary of the location of the CM, the principal inertias, and the

eigenvectors is included.

All CM values in m

All inertias in kg - m 2

After initial design:

Xcm = -.2552 Ycm = -.1196 Zcm = -.3901

Principal Inertias 98000 883

0 0 1 4

I .9995 -030500] Principal Directions -.0305 .9994 0

0 0 1

After 1st iteration:

Xcm = -.1136 Ycm = -.0905 Zcm = -.3909

[1336 0 0 1 Principal Inertias 0 810 0

0 0 1948

Principal Directions the same

After 2nd iteration:

Xcm = -.0571 Ycm = -.0180 Zcm = -.3917

f

k

2085 0 0 1 Principal Inertias 0 845 0

0 0 2733

Principal Directions the same

The final inertia matrix of Cerberus plus propellant is given in Section

2-3 of the report.

The inertia matrices and CM locations were presented here to display

how these quantities changed during the optimization process. The values for

the CM locations are relative to an origin placed in the middle of the thruster

nozzle (See Figure 2-1). The final position of the CM is shown in these figures.

36
Annendix B-2:

Minimization of Truss Mass

t_

The most intense loading on the structure will occur during launch

form Earth. The acceleration of the spacecraft upward inside the launcher

will produce forces well above the normal lg felt by the stationary craft.

Since the 10 bay bus and antenna are off-the shelf items, it was assumed that

they would be able to handle the launch loads without damage. The only

analysis left to do is on the trusses that will be stressed during launch.

The fact that all four booms will be stowed during launch introduces the

necessity for a stress analysis of the supporting truss structure. The

configuration and labels are shown in Figure 2-7. A simple truss analysis

yields the maximum force in the structure [5].

P b

Fmax- a cos 0 (Eq. B2-1)

oq

4

b

\

P

Figure B2-7: Truss configuration during launch

37
P is consideredto be the mass multiplied by the quasistatic load factor. For this

analysis, the load factor was taken to be 10g [4]. The maximum stress in the

structure is Eq. 2.1 divided by the area of the element. This stress must be less

than the yield strength divided by the safety factor. Using this concept and

the geometry of the structure, the mass of an element can be found.

Masselement- 2 sine (Eq. B2-2)

SF = Safety Factor

L -- Length of element

The major finding of Eq. B2-2 is that the the mass of the truss elements

will be minimized by the lowest weight to strength ratio. This assumes

equivalent geometry. For this reason, the truss elements that must carry the

load during launch will be made of titanium. The other truss elements,

especially the members that are retracted in the mag cannister, can be made

of aluminum, since they are lighter and don't need to carry as much structural

load. This will save weight on the overall spacecraftconfiguration.

38
Section 3-1:

Section 3: POWER AND PROPULSION

INTRODUCTION

!"i

In the integration of Cerberus' power and propulsion systems, there

were key requirements that drove each design. The method of attack was to

first identify the requirements for each system. Of secondary importance is

the identification of interfaces with other subsystems of the spacecraft.

Finally, to meet the requirement of feasibility and cost, the power and

propulsion system was discretized into three components: power, propulsion

module and Earth Launch Vehicle (ELV)/ upperstage transfer vehicle.

The power systems main concern was lifetime, since it will take

approximately 19 years to reach Plutonian space. Lifetime of this system will

not preclude it from fulfilling its mission or other possible missions. The

reliability is governed by the space worthiness of the off-the-shelf items used.

By interfacing with all the other subsystems, the peak power usage was

determined for a worst case scenario.

For the propulsion module, the A v's needed for the mission were the

driving factors. Simplicity, reliability and space worthiness, of course, played

an important role in the final design. Tight integration with the mission

planning group produced reasonable Av's that minimized weight and

complexity of the overall design. Masses from all groups were then needed to

get a spacecraft dry weight. This was used to calculate total amount of fuel

needed and final module weight. A final integration with the structure group

was then done.

The driving factor for ELV/upperstage transfer vehicle was the

minimization of on-orbit assembly. It is felt here that for the proposed flyby

no on-orbit assembly should need to be done, thus reducing risk to manpower

and increasing cost savings. Also, the reliability and availability of these

vehicles were critical considerations.

This, again, enhanced the need for reasonable Av's, especially at Earth.

Integration with all groups was then completed in order to fit Cerberus into an

existing or near existing transfer vehicle with no need for on-orbit assembly.

39
i

E

Each sub-system is discussed in detail below. Final consideration in the

method of attack of feasibility and cost is then addressed.

Section 3-2: POWER

For a peak power estimate, a worst condition case was asked of each

subsystem.

could need.

on condition.

spacecraft.

This scenario allowed for the highest level of power each system

These powers were then added in series simulation an all systems

Table3-1 gives these results for each system and the total for the

Table 3-1" Peak power estimates by subsystem

Subs_,stem

Science

(EE 60

AACS 60

Propulsion

TOTAL

Power IW)

30

15

165

For accessories and propulsion this took into account the heating of valves and

catalyst-beds for the system. Structural thermal control was done by separate

units supplying their own heat.

This total number is a conservative power estimate. To allow for error, a 30%

safety factor was added, bringing this estimate to 214.5 Watts. Since the length

of time to Plutonian space is 19 years, it is the necessary to find a power source

that can provide this peak power for the duration of the flight. This will

ensure successful completion of the mission. Also, there is a chance that the

power source will be operational after the rendezvous with Pluto.

The existing tested and flight proven sources are batteries, solar cells,

and Radio Isotope Thermoelectric Generators (RTGs). Of these, use of RTGs is

the only viable option for a mission of this duration. Batteries will not last 19

years and solar cells will be useless at 40 A.U. Other sources are presently

being developed (e.g. heat stirling engines and small nuclear reactors) but are

not yet reliable for long-term use. RTGs, however, have a proven track record.

40
Those employed in Voyager have been operating for over 14 years and have

shown excellent performance[5]. The latest RTGs are built around the General

Mill II I1111

IN %1 II A l I1 tN

Figure 3-1: GPHS/MITG

MII)%PAN )_1 AI

%4111}11 I %IPI'IIq)H ;

A new design that uses the GPHS shown in Figure 3-2 is called the Modular

Isotopic Thermoelectric Generator (MITG) incorporates better features than

the GPHS/RTG[8]. The advancements are due to better design and new

materials that increase the conversion efficiency. These features are listed in

Table 3-2. A typical MITG "slice", shown in Figure 3-3, produces approximately

24 Watts at 28 Volts with an initial thermal load of 250 Watts.

al

:E':"_;_ _ :: r

OF POOR Qt.IALLm_,'

Figure 3-2: Showing the MITG unit

41
.'t

_i if-X,

OF POOR _i I_: "',',_,

Figure3-3: Showing a typical MITG slice

Table 3-2: Advancements for MITG

2

4

Adaptable to wide range of power since it is

available in standard 24 Watt slices

When varying number of slices, only the

redesign of housing is necessary

Performance of each slice can be checked

individually

The series parallel circuit permits high

redundancy

Lighter weight (higher specific power)

A comparison was done by the producer of the MITG (Fairchild Space

and Electronic Company). between a typical 290 Watt GPHS/RTG and a similar

42
282 Watt 12 slice MITG[8].

3.

The results of weight differences is given in Table 3-

Table 3-3" Showing weight advantage of MITG over a typical RTG

_qq, l_ _'. _qOd " ,,_ aWIG5iCe_P. lWl

Im

m 1Mml_

_C.km

_ evta_t

I_m,e s_m

ewt_ ks_

_ mmw

Im4

r_

i iEii!ii!

EE

te

leeM_ _

tN | iw, m lul_

ii Ilj _ llk4 kT BIle

The figures show that for approximately the same power output the MITG

design weighs about half as much. This weight savings relates to a cost

savings while still using a reliable and proven GPHS, therefore satisfying

mission requirements. For Cerberus, two 11- slice MITGs are slated to be. used.

Each MITG is capable of providing the spacecraft with the necessary power,

thus achieving 100% redundancy. As mentioned before, a 30% safety factor

has been incorporated. This has been done for sources of errors in predicting

performance such as effect of fuel decay on power transfer, uncertainty in

the amount of dopant precipitation in the thermal electric material, loss due to

oxygen diffusion in 238puO2 pellets, and uncertainty in establishing power

profile throughout the mission [2]. For 11 slices, Figure 3-4 shows that an

optimum can be found at the elbow of this curve. At MITG weight of 24.95 kg

(551b), the resulting power supply for Cerberus is then given in Table 3-4. The

equations for the calculations are shown in Appendix C.

43
f

i i i i : ....11.......

Lk

./ .... :...:._.:i...,..'." 2..:.,.:; ..,--:i.. ..., .".) ..

_............ ../...-......._.:.. _ ._,,.. : ; • " .y" 12 _ .

Figure 3-4: Performance curves for MITG

At

Loading

After 19

Years I

Table 3-4: Cerberus Power Supply (MITG)

Thermal Load/MITG

tWattsl

2750

2369

Output/MITG

/Wattst

259

222

% Above Peak

IWatts)

57

35

Specific power 10.36 Watts/kg. Total weight = 49.9kg

NOTE: The 11 slice system saves a total of 2.21 kg/MITG over the 12 slice system

while easily satisfying power requirements

It can be seen that the 11 slice MITG slightly exceeds the 30% safety factor.

If exact power was desired, the MITGs could be fined tuned simply by adjusting

the radiator fin lengths. The power conditioner along with the computer will

regulate and condition power according to the needs of the spacecraft. This

will be done by autonomous sensing and programming that will periodically

review the system.

In conclusion, with the conservative approach taken, it is felt that this

is a feasible and cost effective system due to its savings in weight and use of

off-shelf items. If necessary, this system could easily be up or down-sized for

the possibility of other missions such as the measurement of heliopause or a

change in launch date.

OF POOR QGh_I. ITT'

44
7-'

Section 3-3: PROPULSION MODULE

After the spacecraft is delivered on to the transfer orbit, the propulsion

module's function is to carry out the necessary Av sequence up to and

including the Jupiter assisted flyby. It must also carry of all of the necessary

fuel for both propulsion and AACS. The Av sequence is listed below in Table 3-

5.

Table 3-5: Cerberus Av sequence

Location

Earth

i

Av(km/s 1

.001

Midcourse .395

Jupiter .391

Unmanned reconnaissance spacecraft of the past, such as Voyager,

have weighed between 200-800 kg [4]. This is the weight class that Cerberus

was designed for. Table 3-6 show Cerberus' dry weight breakdown. For this

payload mass and the required Av's, the fuel of choice is Hydrazine (N2H4). It

has an excellent track record with 20 years experience and a large data base

[7]. It's Isp of 235s provides adequate thrust times for this type of spacecraft.

Table 3-6: Cerberus' dry weight breakdown

Payload Type

Science

Mass (k_[ t

90

MITG'S 50

Antenna 40

Propulsion Module 90

AACS 85

iO3U

Structure

iTOTAL

35

250

640

45
f

Other fuels were considered , such as cold gas and the bipropellant the

N204/MMH. Cold gas offered a simpler design and is less expensive but it's Isp

of 50 is only good for low thrust pulsing. The bipropellant has a higher Isp of

285 but needs a more complex system of metering the fuel. For a mission of

this lifetime the simpler the system is more reliable. Therefore, the Cerberus

propulsion system will be based around the monopropellant hydrazine. A

more complete breakdown of hydrazine's advantages are shown below in Table

3-7 and were found in [7].

Table 3-7. Hydrazine advantages for this mission

1

2

Simple and reliable (20 years experience)

Lowest cost propulsion system, other than

cold gas

Space storable for long periods (> 12 years

demonstrated

4

5

Low thrust capability

Moderate thrust levels

To get the amount of fuel needed a semi-dry mass was worked backward

using the rocket equation. By semi-dry mass it is meant that part of the AACS

fuel will still be left after the Jupiter flyby. A 20% redundancy was then built

in to the calculations. The equations are shown in Appendix C.

These calculations yielded the propellant structure weight along with

the amount of propellant needed. The total amount of N2H4 needed, including

the AACS requirements and a 20% redundancy, is shown below.

392

+81

Propulsion for z_v's + 20%

Attitude and control (ACCS)

TOTAL = 473 kg

46
t •

i

A two tank configuration was then chosen. Two rubber bladders carry

the hydrazine and two smaller helium pressurant tanks are needed to pressure

feed the propellant on demand to thrusters. This configuration was

determined to be better than a 3 or 4 tank set up because of its simpler design.

Although simpler, redundancy was still achieved.

The size of the tanks or rubber bladders was done through a simple

conversion of propellant mass to volume through the density of hydrazine.

The radius of the tank was found by equating this quantity to the equation for

the volume of a sphere. As an approximation, the helium tanks were taken to

have 1/3 the radius of the rubber bladders.

A 400 N main engine was selected to deliver the Av's. Rough

conservative calculations of the thrust times needed to perform each A v are

given in Table 3-6. The equations and an explanation of the estimates are

given in Appendix C. These times are still feasible for approximate impulsive

maneuvers. Furthermore, such engines have been used in past and are

present missions. They have proved themselves to be reliable in situations

with similar Av's [4].

Table 3-6: Times of thrust for Av's using 400N engine

Avlkm/s 1

.001

Time

2.3sec

.395 17min 59sec

.391 14min 39sec

The complete design of Cerberus' propulsion module is given in Figure

3-5 and is shown in 1/25th scale. The final integration into the overall

spacecraft structure is shown in Figures 2-1 and 2-2. The propulsion module is

shown fitting into the 1.5m wide main bus of the structure. The side view

shows that half the module will be up in the bus. The system configuration

and valve network is shown in Figure 3-6. Since thrusters will not be

redundant, 2 valves and 2 lines are assigned to each thruster providing a feed

redundancy from each tank.

47
r  ̧

Table 3-7 provides the weight breakdown for the final propulsion

module. It should be noted that the weights for components are rough

estimates based on findings in reference [6]. While the weights are not exact

there is room for expansion of the overall design if needed. Also, the weights

for AACS thrusters and plumbing were not taken into consideration here but

instead are accounted for in the AACS weight of 85kg.

48
Hydrazine

Bladders

Helium Tanks

Propulsion

Module and

supporting

structure

1/25th Scale

4cm=Im

Figure 3-5: Propulsion Module

Top and Side View

49
Relief

Ill

:.: x :.x,_ I

Fill

1

Regu

i

1

ator

_1

II

Regu

Fill

I1

ator

!

[]

IX,X

i111111 ] I !i il Iil c

/ / x

9N ACCS

Thrusters 400N

Thruster

II

xiI

Regulators

Relief

N

i|

Figure 3-6. System configuratlon

5o
Table 3-7: Dry weight breakdown for propulsion module

Component

Tanks (2 bladders, 2 He tanks)

PMDS Management devices

400 N engine

Heaters

Weil_ht(kg)

27

5

Structure 44

Residuals 1

Iror 9o

i The cost of the system has been kept to a minimum by its light weight

and simple design. In addition, the propellent to be used, hydrazine, has been

flight-tested and is reliable. With a redundancy in the fuel (20%) and in valve

configuration, this module will last the the mission lifetime while not

precluding it from being utilized for other missions.

Section 3-4: ELV/UPPERSTAGE TRANSFER VEHICLE

Once the final propulsion sizing was finished, the fully loaded (wet

weight) of the spacecraft was determined. This was part of the calculation for

the propellant need done in the Appendix. The wet weight of Cerberus craft is

1093 kg. The Av needed to insert Cerberus into its transfer orbit is 5.192 km/s.

Since most upperstages utilize the solid propellant ammonium perchlorate, it

was used for the calculations of amount of propellant needed to do this burn

with this payload. The burn required 6401 kg of fuel making the total weight

of the upperstage 6957 kg. This number plus the wet weight of the craft

determined the total integrated takeoff payload for the ELV. This came to 8050

kg. Therefore, the upperstage must have 6401 kg of fuel and the ELV must be

able to lift 8050 kg into Low Earth Orbit (LEO).

51
¢.-

(

3

r

r

Table 3-8 shows the available upperstages while Table 3-9 shows the

available ELVS that can meet these requirements[9]. It is noted also in the

tables the approximate percentage of downloading that needs to be done for

each vehicle.

Table 3-8: Possible upperstages

Vehicle

IUS

q-OS

TOS/AMS

Contractor

Boeinl_

Orbital Sciences

Orbital Sciences

Wei_[ht

14_660

10,894

16,016

% Download

57

36

57

Table 3-9: Possible ELV vehicles

Vehicle

Commercial

Titan

Titan 4

NUS (Typel)

Titan 4

NUS (Type2)

Contractor

Martin

Marrietta

Martin

Marrietta

Martin

Marrietta

Performance

to orbit _k_t

14_519l

17,7401

17,015

Approximate

% Download

45

55

53

The tables show that the optimal ELV/upperstage combination would be

the Commercial Titan with the TOS upperstage. The IUS upperstage was

eliminated because it could not be downloaded the necessary amount [3]. Others

were then considered for the least amount of downloading, i.e., the least

alteration to the existing vehicle• Vehicles requiring the least amount of

downloading will cost less and be easier to get flight ready. Again the

requirements were met for the cost effectiveness by the use of off-the-shelf

items.

Section 3-5: CONCLUSION

Through the use of off-the-shelf items and a conservative cost effective

approach, feasible power and propulsion systems were conceptualized for the

52
j -

Cerberus spacecraft. It is believed that these systems will not preclude

Cerberus from successfully completing its mission along with possibly

performing others. It is also believed, from a cost standpoint, that these

systems will not cause Cerberus to overshadow other missions of the same era.

References

1. Campbell, R. W., "Space nuclear power safety from a user's

viewpoint", Space Nuclear Power Systems, 1984, p. 515.

, Chmielewski, A.B., "Improved techniques for predicting

spacecraft power", IECEC, 22nd Intersociety Energy Conversion

Engineering Conference, August 10-14, 1987, p.392.

.

.

Cole, "Phase A report preliminary assessment of space transportation

alternatives for planetary missions," Launch vehicle alternative,

performance, availability and schedule. JPL D-3332 JPL NASA

Document, May 15, 1986.

Garrison, P.W., "Advanced propulsion for future spacecraft", AIAA

81-1534R, AIAA/SAE/ASME 17th Joint Propulsion Conference,

Colorado Springs, CO.,July 27-29, 1981.

. Iserland, K., "Overcoming the launch crisis: A challenge for ELVs",

ACTA Astronautica, Vol. 20, pp. 9-19, 1989.

. Koelle, D. E., et al., "Cost reduction for future communication

satellites by a standardized propulsion module (SPM)", ACTA,

Astronautica, Vol. 17, No. 4, pp. 397-405, 1988.

. Koepke, A., "AAE 241 Class notes", University of Illinois, Kinkos

216.05, 1990.

Packet

.

.

Schock, A., "Modular Isotopic Thermoelectric Generator", Fairchild

Space and Electronics Company, Germantown, MA. 20769, 1980.

No author, "Aerospace Forecast and Inventory"

Aviation Week and Space Technology, pp163-173, March 19, 1990.

53
Annendix C: EOUATIONS

EOUATIONS FOR POWER CALCULATIONS

Using the power law equation:

i

p=ce kt

and using conditions found in ref.[2] for 238puO2

eq. 1

1.5493 years later

t(0)= 4460.6 watts

t(1.5493)=4406.7 watts

the constant (k) for 238puO2 was found to be -7.8468E-3

Now plugging the conditions of c=initial thermal loading of

258.5 watts from Figure 3-4 for an eleven slice MITG and t=19, the power

after 19 years is 222.69 watts.

EOUATIONS FOR PROPELLANT CALCULATIONS

Using the rocket equation:

A v=(Isp* g)*ln(minit/m final) eq.2

and the propellant mass fraction:

mprop mass frac=

mprop struc.+prop

eq.3

masses were found.

For Hydrazine: Isp=235 mass frac.=.9

For Ammonium Perchlorate (70%)'

unusable=2.5%

Isp=2 mass frac.=.92 unusable=2%

54
EOUATION_ FOR CALCULATIONS OF TIME OF BURNS

A simple linear approach was taken using the following:

F=ma eq.4

t

to get the acceleration (a) and the linear velocity equation:

Av=at eq.5

The change in mass was accounted for at the end of each burn,

then subtracted from old to get new mass for next burn.

*It is noted here, that in reality (a) is not constant throughout the burn

since the mass is also changing as fuel is expelled from

the spacecraft. Iterating numerically would yield smaller times

since the rocket effect would occur. Therefore, this linear

approach is a rough but conservative estimate.

i •

55
r

Section 4: ATTITUDE. ARTICULATION. AND CONTROL (AACS)

Section 4-1: INTRODUCTION

The function of the Attitude and Articulation Control Subsystem (AACS)

is to determine the orientation of the spacecraft and control its motion. This

includes orienting the axes of the spacecraft; controlling the valves, heater,

and firing of the thrusters; firing the engine for trajectory correction

maneuvers (TCM); and controlling the science platform [1].

The design of the AACS was determined largely by the requirements in

the Request for Proposal (RFP). The requirements that applied specifically to

the design of this subsystem were satisfied. First, it was required that the

spacecraft's performance, weight, and cost were optimized. Second, the

spacecraft was designed to be simple, reliable, and easy to operate. Third, offthe-shelf hardware and technology available by 1999 were used as much as

possible. Fourth, the spacecraft was designed to be able to perform several

possible missions. Fifth, the spacecraft will have a design lifetime sufficient to

carry out its eighteen year mission plus a reasonable safety margin. (A 20%

safety margin would result in a design lifetime of 21.6 years.)

In addition to those in the RFP, there were design requirements dictated

by the other subsystems. The Command, Control, and Communication (CCC)

Subsystem required that the high-gain antenna must be pointed at the Earth

with 0.1 ° pointing accuracy. The Science Subsystem needed to be able to point

remote sensing instruments at specific locations for extended periods of time

with 0.1 ° pointing accuracy. The Mission Planning Subsystem required that

windows must be identified in which the AV maneuvers could be executed, and

the Power and Propulsion Subsystem defined limits for the power consumption

and mass of the AACS.

Finally, the AACS was also designed to follow the overall objective of the

Cerberus mission. Cerberus was designed to sell and work, which means that at

every turn, measures were taken to design an AACS that was cost effective and

truly feasible.

£

56
Section 4-2: MAJOR FEATURES OF THE AACS

After careful consideration between three-axis, spin, and dual-spin

control, three axis stabilization was selected as the control method for

Cerberus. Spin stabilization met many of the requirements. It is the simplest

and least expensive of the three [2]. It is very reliable and has a long lifetime,

which is very important considering the length of this mission (eighteen

years). In addition, it is the lightest and requires the least power [3].

Nonetheless, spin stabilization was unacceptable for one reason. The Cerberus

mission requires that the remote sensing science instruments be inertially

fixed for extended periods of time. This would require the spacecraft to

undergo a complicated despinning process, which makes spin control

infeasible. On the other hand, both three-axis and dual-spin provide the

neccesary fixed inertial orientation. Dual-spin has certain advantages over

three-axis. The former provides scanning science capabilities and has low

sensitivity to disturbances, whereas the latter has neither [3]. The deciding

factor between the two, however, was the fact that dual-spin is much more

expensive and complex than three-axis. Although dual-spin offered certain

conveniences, it did not meet the requirements of optimizing cost and

maintaining simplicity. Therefore three-axis stabilization was selected. Table

4.1 compares the advantages and disadvantages of the three control methods

discussed above.

Table 4-1: Comparison of Control Methods

rypes of Control

Three-Axis

Spin

[2, 3]

Advantages

,High accuracy

,Good maneuverability

,Adaptable to changing

mission requirements

Allows inertial remote

sensing science

,Simple, low cost

,High reliability, long life

,Low weight and power

Inherent science scan motion

,Low sensitivity to disturbances

Disadvantages

• High weight and power

• Costly hardware

•Extensive fault detection/correction for backup

• Poor maneuverability

• Must despin to do some

imaging science, which is

complicated process

57
Dual-spin

Provides both scanning _Expensive and complex

and inertial science ].Articulated elements reLow sensitivity to distur- | quire balance compensabances / tion Fixed inertial orientation

After the decision was made to use three-axis control, the next step was

to decide how to best implement such control. There were two options. The

first was to use momentum wheels or control moment gyros for stability and

small turns. A reaction control system using gas thrusters would also be

needed to dump momentum from the momentum wheels or control moment

gyros and to do large turns and maneuvers. The second option was to only use

gas thrusters. A trade study was conductedto see if the addition of momentum

wheels or control moment gyros would save enough fuel to offset the

additional weight. First, it was concluded that momentum wheels are lighter

than control moment gyros [2], so momentum wheels were used as the basis of

comparison. From three different sources[2, 4, 5], the mass of a system of four

momentum wheels was estimated at 100 kg. Next, the mass of the fuel that

would be saved by using momentumwheels was calculated to be 80.6 kg. See

equation (1) in Appendix D for details of this calculation. Therefore, even

without considering the additional fuel needed for momentum dumping, using

a system with only thrusters would be lighter than one that used both

thrusters and momentum wheels. See Table 4.2 for a summary of this trade

study.

Table 4-2: Trade Between Momentum Wheels and Thrusters

System with both momentum

wheels and thrusters

System with thrusters only

mass (kg) 80.6 100.0"

*Does not include fuel needed for momentum dumping

In addition to being heavier, momentum wheels would also add unnecessary

complexity and would decreasethe lifetime due to wear. Therefore, three-axis

control will be implemented using a configuration of only thrusters because

such a selection satisfies the requirements of simplicity, sufficient lifetime,

and low weight.

58
The last major feature in the design configuration of the AACS is the

articulation control of the science scan platforms. The Science Subsystem

requires both a high accuracy scan platform with two degrees of freedom and

a low accuracy platform with one degree of freedom. Two options were

explored. The first was the traditional method, in which scan platforms are

articulated using two step motor actuators without any momentum compensation [6]. In addition, the AACS electronics, star tracker, and gyros are all

placed on the spacecraftitself. This was the technology used for Voyager. The

second option involves a new technology being developed for the Mariner

Mark II (MMII) project, and it is called the Integrated Platform Pointing and

Attitude Control Subsystem(IPPACS). If the second option were used, the star

tracker, gyros, and AACS electronics would all be placed on the scan platform

[7]. This second option using IPPACS was selected for the following reasons.

First, the scan platform momentum compensation decouples the scan platform

dynamics from those of the spacecraft [7]. Second, this decoupling of the

dynamics ensures dynamic stability of the spacecraft [8]. This satisfies the

requirements of reliability and ease of operation. Third, the high accuracy

sensors and controls are rigidly attached to the high accuracy science

instruments, which greatly reduces many errors found in systems like that on

Voyager [7]. See Table 4.3 for a comparison of errors between Voyager and

MMII IPPACS. This meets the requirement to optimize the spacecraft's

performance.

Table 4-3: Comparisonof Errors BetweenVoyager and MMII IPPACS in Terms

of Scan Platform Pointing Control (Adapted from [71)*

Limit Cycle

Error Source

Sun Sensor

Star Sensor

Scan Platform Control

Structural Misalignment

Dynamic Stability

Gyro Drift (2 hours)

Total 3c Scan Platform Pointing Accurac Y

*All numbers in mrad

0.69

0.59

1.20

MMII IPPACS

0.00

0.00

0.02

0.34

0.58 0.58

0.33 0.33

0.00 0.11

2.26 0.76

59
0 -

Because IPPACS has such high accuracy, it will easily be able to perform

several possible missions. Although IPPACS has yet to be flight tested, it will

have flown aboard a Mariner Mark II mission before 1999, and can be

considered off-the-shelf hardware. The biggest disadvantage of IPPACS is that

it shifts the weight of many of the attitude control components away from the

center of mass to the scan platform, slightly increasing the moments of inertia

of the spacecraft [7]. Nonetheless, the advantages far outweigh this disadvantage, and IPPACS was selected to control the articulation of the high accuracy

platform. Figure 4.1 shows a schematic of the IPPACS configuration, and

Figure 4.2 shows the actuator orientation for two articulation degrees of

freedom.

The question still remained of how to control the articulation of the low

accuracy platform. One option was to use IPPACS on it also, and therefore have

a fully redundant AACS. This, however, would certainly not optimize weight

and cost. Instead, it was decided to use a single uncompensated step motor

actuator and measure the angular displacement with a simple optical sensor.

Momentum compensation wasn't necessary because the low accuracy platform

will not be used for determining the attitude of the spacecraft or conducting

imaging science. Instead, it will be used merely to scan free space. The

absence of momentum compensation will, however, lead to some disturbance of

the spacecraft which must be corrected by firing the thrusters. This will still

be lighter and less expensive than including a momentum compensation

wheel and thus meets the RFP requirements of optimizing cost and weight.

6O
jr.-

CONSTIU_IN1"$

• _OCK F;X_ ;N $fC

• CONEJ. TOCLOCK

• I. J. tAXES FIXEO

IN PLATFORM

• ALLAXESTHROUGH

PLATFORMC.M.

• Ijj • Ikl_

Figure 4-1: IPPACS Configuration (Adapted from [8])

WH£E). r'_

STEPMOTOR

Figure 4.2" Actuator Orientation for Two Articulation Degrees of Freedom

(Adapted from [8])

61
i

f

Section 4-_-" HARDWARE SELECTION AND PLACEMENT

After the decision was made to use thrusters for attitude control, the

next phase was to select thruster size and placement. With the use of IPPACS

technology, turning rates of less than 1.3 deg are desired. From preliminary

sec

data on the moments of inertia and thruster lever arm distances, 0.9 N

thrusters were found to produce a turning rate of 1.261 (leg which meets the

SeC '

specification. See equation (2) of the Appendix. The 0.9 N thrusters were

selected because in addition to meeting turning rate specifications, they are

off-the-shelf hardware, being previously used on both Voyager [9] and MMII

[7]. Hydrazine was selected as the fuel so that these thrusters and the 400 N

engine could be supplied by the same fuel source. This satisfies the

requirement of simplicity.

The thrusters were placed in such a way as to provide a torque couple

around each axis in both directions of spin. This is required for three-axis

control. Such a configuration requires four thrusters per axis, or twelve

thrusters total. Figure 4.3 illustrates the placement of the thrusters with

respect to the principal axes of the spacecraft. Only the minimum number of

thrusters necessary was used for this configuration, which appears to

sacrifice reliability and redundancy for the sake of optimizing weight and

cost. This, however, is not actually the case. The thrusters themselves have

proved very reliable on past missions; most problems occur in the plumbing,

valves, and heaters which are redundant on Cerberus. The details of this can

be found in the Power and Propulsion Subsystem in Section 3. Also, the

thrusters were carefully placed to ensure that

62
! •

{

t

r

k

÷ z (RolO

\I/

/I\

+X (Pitch)

\

/\

/

/

\1/

/1\

\1/

+Y (Yaw)

/1\

Figure 4-3: Placement of the Thrusters With Respect to the Principal Axes of

the Spacecraft

their exhaust would not damage any of the science instruments. Refer to the

Structures Subsystem in Section 2 for a picture showing thruster locations

relative to the scan platforms.

The Planetary ASTROS (Advanced Star/Target Reference Optical Sensor)

was selected as the star tracker for Cerberus. Its accuracy of 4 arcsec (0.001 °)

exceeds the required pointing accuracy of 0.1 ° [7]. The Planetary ASTROS has

the additional benefit of being able to track other targets, so it will also be used

to locate Pluto and Charon. In addition, it has the lowest weight and power

requirements of any star tracker currently available (the Voyager Canopus

Tracker is no longer manufactured) [7]. There is, however, one problem area.

Although the Planetary ASTROS is internally redundant and has the longest

lifetime of any star tracker currently available, its design lifetime of ten years

falls dreadfully short of the required 21.6 years [7]. Nonetheless, the best plan

of attack is to select the Planetary ASTROS and continue to work on increasing

its lifetime through design modification. A comparison of the characteristics

63
of the Planetary ASTROS, Canopus Tracker, and the Digital Standard Star

Tracker (DSST) are given in Table 4.4.

Table 4-4:

(Adapted from [7]

Canopus

Comparison of Star Tracker Characteristics

Parameter

'Field-of-View (degrees /

_Spin (drift_ Rate (deg/sec)

Calibrated Accuracy (arcsec)

Multistar Measurement

Internal Redundancy

Tracker

9x36

N/A

180

DSST

8x8

<0.3

Planetary ASTROS

llxll

< 0.5

Mass (k_q)

Power (W)

10 4

NO NO 3 stars simultaneously

NO NO

8.2

20.0

4.3

4.5

YES

8

11

The Planetary ASTROS will be placed on the high accuracy scan

platform as necessitated by IPPACS. It will be boresighted with the science

instruments to enable accurate pointing of these instruments. Furthermore,

placement of the star tracker on the high accuracy scan platform allows the

Planetary ASTROS to track stars and other objects without requiring the

rotation of the entire spacecraft. This will save AACS fuel and optimize weight

and cost. Because the Planetary ASTROS is already being developed for the

MMII, Cerberus will reap the benefit of a recurring cost as opposed to

spending money to develop a new technology. In addition to meeting the

requirement of optimizing performance, weight, and cost, the Planetary

ASTROS also qualifies as off-the-shelf hardware.

In selecting the components for the Inertial Reference Unit (IRU),

there were found to be two general choices: flight-tested mechanical gyros or

a new gyro based on fiber optics called the Fiber Optic Rotation Sensor (FORS).

One source states, "FORS is attractive for space missions because it promises

performance comparable to or better than that of mechanical gyros as well as

significant improvements in lifetime, weight, power consumption and cost"

[10]. As with the Planetary ASTROS star tracker, the FORS lacks flight

experience and a sufficient lifetime. Nonetheless, it is still the best available,

having a longer lifetime than any mechanical gyro because there are no

moving parts in the FORS [7]. Clearly, an optical gyro such as FORS best meets

the requirements. The question remained, however, of which optical gyro was

best. Of the three optical gyros studied, FORS best optimized performance

64
(lowest drift and rate noise, and best angular resolution), lifetime, power, and

weight. Table 4.5 gives the numerical data for these parameters.

Table 4-5: Optical Gyro CharacteristicstAdapted from [7])

DRIRU II CG-1300 Laser Gyro

FORS

0.2E-3

Parameter

Residual Drift Rate (deg/hr)

Rate Noi'se (deg/sec)

7.0E-3

1.0E-5 40E-5

Angular Resolution (arcsec) 0.005 0.05 1.4

10 3 3

< 10"

10

MTBF (yr)

Power (W)

Mass (kg)

Volume (in 3)

*For three units

22

11

990

18

18

1000 350

In addition to meeting the above stated requirements, the FORS will be off-theshelf before 1999 because it is tentatively scheduled for use on the MMII

project. For these reasons, Cerberus will use FORS for its IRU. Two identical

sets of three (one for each axis) will be placed on the high accuracy platform,

making the IRU redundant [7]. This is to fulfill the requirement of reliability.

A sun sensor will be mounted behind the high-gain antenna and will be

pointed at the sun through a hole in the reflector. As with Voyager, it will be

boresighted at an offset of 5-6 ° off the Z (roll) axis of the spacecraft, so that

when the sun sensor points at the sun, the antenna points at the Earth, as

required by CCC [6].

At this point in the preliminary design, there was no way to calculate

the precise A V required for a minimum maneuver scheme. Instead, a

calculation was performed based on an average of AACS maneuvers per month

in past missions. This calculation is given in equation (1) of the Appendix.

The result is that 80.6 kg of hydrazine will be required for AACS maneuvers.

This does not include TCM's.

To conclude the discussion on hardware selection and placement, Table

4.6 gives a summary of the power and mass requirements for the AACS.

65
,i

Table 4-6: Power and Mass Requirements for the AACS

Component

Planetar), ASTROS Star Tracker

FORS IRU (3-axis r redundant)

,Digital Sun Sensor

[Microstep Scan Actuator

Low Accuracy Platform Actuator

Low Accuracy Platform Position Sensor

Heating valves and thrusters

Total*

2

Power (W)

11

10

19

Mass

Ik_)

8

10

4

21

1 7 6

1 1 0.5

12

*Does not include computer, propulsion system, or fuel

7

60

15

64.5

Section 4-4: SCANNING AND POINTING REOUIREMENTS

IMPLEMENTATION

By pointing the sun sensor at the sun and the star tracker at a star

orthogonal to the roll axis, the high-gain antenna can be pointed at the Earth

with 0.1 ° accuracy as required by CCC. From this orientation, the low accuracy

scan platform can perform particle field science around its one degree of

freedom; the Science Subsystem did not deem it necessary to perform such

experimentation around all three axes. From this same celestial lock

orientation, much of the science on the high accuracy scan platform can be

accomplished. The star tracker will be used in this situation to track the

targets of the science instruments. If an object cannot be sighted from

celestial lock, the entire spacecraft will be rotated using the 0.9 N thrusters,

with the attitude controlled by the IRU. With this configuration, the remote

sensing and imaging science can remain fixed on a target in any direction for

extended periods of time (approximately three hours) as required by the

Science Subsystem.

Section 4-5: ATTITUDE CONTROL MODES

After Cerberus separates from the launch vehicle, the AACS will enter

the deployment mode. During this time the AACS will control the deployment

of the magnetometer, RTG's, science scan platforms, and high-gain antenna

66
i"

[9]. All the hardware and instruments will be checked out and calibrated in

order to optimize performance.

When the checkout and calibration is complete, Cerberus will enter the

cruise mode. During this mode, the AACS will execute a preset series of

commands to measure fields and particles of interplanetary space [9] after

every A.U., as required by the Science Subsystem. Other procedures executed

during this mode are attitude determination, high-gain antenna pointing [1],

and maintenance of all three axes within a deadband limit of 0.05 ° [6]. This

mode would also include a special routine in both the computer RAM and ROM

to point the antenna back to the Earth from any other orientation. This

enhances the spacecraft's reliability and ease of operation.

Finally, Cerberus will enter its flyby mode when it encounters Earth,

Jupiter, and finally Pluto. At Earth and Jupiter this will involve the AACS

recognizing a box-shaped window defined by certain stars in which it will

control the firing of the 400 N engine. During the approximately fifty day

encounter near Jupiter and Pluto, the AACS will control the pointing and

slewing of the high accuracy scan platform.

Section 4-6: CONCLUSION

The AACS of Cerberus will use a three-axis stabilized design controlled

with 0.9 N hydrazine thrusters. The high accuracy scan platform will be

controlled using the new IPPACS concept which offers the tightest pointing

and the most reliable control. All components were selected to optimize

performance, weight, and cost and will be readily available by 1999. The AACS

has also been shown to meet the requirements imposed upon it by the other

subsystems. The critical problem areas left to be solved are extending the

lifetime of the components to at least 21 years (20% safety margin) and designing the details of the attitude control modes and the exact sequence of

AACS commands.

67
References

1. Galileo: Exploration of Jupiter's System, NASA SP-479.

2. Wertz, James, ed. Spacecraft Attitude Determination and Control.

3. Gates, Robert L. AACS lecture. {AAE 241 class notes on 3/22/90}.

. Kendig, John. R. and Davis, Porter. "Space-qualified 1000 Foot pound-second

Magnetically Suspended Reaction Wheel Assembly (MSRWA)." Guidance

and Control 1981, Vol. 45 of Advances in the Astronautical Sciences.

5. Research and Technology 1986. NASA Goddard Space Flight Center.

6. Voyager Neptune/Interstellar Mission.

Mission Systems Handbook.

Flight Science Office, Science and

7. Draper, Ronald F. The Mariner Mark II Program, AIAA-84-0214.

Propulsion Laboratory, Pasadena, CA.

Jet

. Bell, C. E. and Stanton, R. H. InCr_ially Referenced Instrument Pointin_

Platform with Moment_lm Compensated Articulation. Proceedings of the

SPIE National Symposium and Workshop on Optical Platforms, Vol. 493.

9. Adamski, Terrence P.

AIAA-87-0501.

Command and Control of the Voyager Spacecraft,

10. De Paula, R. P., et al. Fiber Optic Rolation Sensor for Space Missions.

Proceedings of SPIE, Vol. 838, Fiber Optic and Laser Sensors V (1987).

68
aammaLr_-12

m l = mass for small AACS maneuvers

= 0.373 kg {from [6]}

month

masstot = .373month_. 1 year 0 kg y12 monthS'_l 8 years) f

= 80.6 kg eqn. (1)

2_ Ct?

0c =-i"_-v t_" ) {from [3]}

eqn. (2)

69
?

Section 5: COMMAND. CONTROL AND COMMUNICATION

Section 5-1: INTRODUCTION

For the Cerberus mission, communication between the spacecraft and

Earth becomes the primary driver for the Command, Control and

Communication (C 3) subsystem. Due to a time lag as high as 11 hours (roundtrip), the craft has been programmed for a high degree of autonomy. Using as

much off-the-shelf hardware as possible, it will use the latest in computer

technology, and utilize a spare 4.8m Galileo antenna. The receiving antenna

system will be the existing Deep Space Network (DSN). Decisions were reached

based on previous missions and the requirements specific to the proposal.

Section 5-2: COMMAND AND CONTROL

Consisting primarily of the computer, Command is divided into three

main subsystems Computer Command Subsystem (CCS), Attitude And

Articulation Control Subsystem (AACS), and Flight Data Subsystem (FDS). Each

subsystem will utilize expert systems when available, as well as sophisticated

fault protection algorithms, (FPA). Due to the distance involved and the length

of the mission (18.7 years) the spacecraft is designed to be almost fully

autonomous.

Command will be performed by two computers in parallel redundancy.

Each computer will consist of five central processing units (cpu) with math

coprocessors, all accessing a common memory chunk of 64 megabytes (MB).

One cpu designated as the monitor will break each assignment into subparts

and allocate them to the other four cpu's. The monitor will also serve as a fault

detector should one of the cpu's fail. In this case, an FPA will be used to

exclude that cpu. In the event that an entire computer fails, the cpu's in the

remaining computer can be reprogrammed so that they are working in

redundant pairs. In a worst case scenario, the whole spacecraft could be run

from one cpu.[17]

The maximum bit-rate for Cerberus is 145.5 kbps. Science

instrumentation may require up to 1000 kbps. Real-time transmission, then, is

not always feasible. Data must be stored until transmission is possible.

External memory is needed.

70
,t

f

Three options exist for external memory magnetic tape, removable

hard drives or optical disks.J6] Although not proven as space technology,

optical disks provide a compact, cost efficient, error-free method for storing

data. It will have read-write capabilities, with a maximum storage capacity of

600 MB. See Table 5-1.

Table 5-1" Comparison of External Memory Options

External Memor]¢

Magnetic Tape

Removable Hard drive

Optical Disk

Disadvantages

Bulky, error-prone,

heavy.

Not proven tech., high

power cost_ heavy.

Not proven tech.

Advantages

Proven technology,

inexpensive.

Small, low error rate.

Small, lightweight, low

error rate.

Three options also exist for the computer programming language.

Assembly is the lowest level language of the three, and as such is difficult to

program in. C is an industry standard. Many people are familiar with it and

programming is much simpler than with Assembly. For this mission, ADA was

selected. The standard for the Defense Department, it is ideal for situations

where prioritizing is needed. Semaphores, which act as flags, signal

important incoming information, allowing new information or more

important processes to take precedence over existing tasks. This is ideal for

sequencing.[ 17]

Each computer will be able to access the optical drive. The computers

and the optical drive will be fully independent and redundant, in case of

failure. In this way, a computer could be reprogrammed from either the

optical drive or the other functional computer. The computers and optical disk

will be redundantly programmed for the three subsystems CCS, AACS, and

FDS.[1]

71
Section 5-2.1: COMPUTER COMMAND SUBSYSTEM

As overseer of the craft, CCS is primarily concerned with issuing

commands to the other subsystems. Other responsibilities include sequencing,

FPA's and real-time command processing.

Sequencing will function throughout the duration of the mission, but

becomes more important when science instrumentation is active. It assigns

each function a value based on priority. This priority is determined before the

mission's onset. Considerations include, but are not limited to, whether science

cannot be achieved from Earth, capability of the instrument(s) required, the

number of observations needed, duration and time of the observations, power

needed, and tolerance of location and duration. A program such as SEQTRAN

will be implemented.[22] This program converts input into mnemonic

commands and checks for constraint violations. Two constraints are

spacecraft physical limitations and mission rules established for safe

operation of the spacecraft. Maximum efficiency of sequencing space is

achieved by overlap. This allows one sequence to start before a prior sequence

is finished.

Expert systems will be used wherever possible. These allow on-line

changes to be made in the structure of craft operations without input from

ground control. Each system must consist of knowledge representation,

knowledge utilization, and a computational model.J29] These will be further

broken down into a data handler and trend analyzer.J7] The data handler

receives data from telemetry and stores it appropriately, according to

significance at that particular time. The trend analyzer calculates, plots and

posts trend information. Use of expert systems will provide better fault

protection, allowing quick response to potential problems.

Improved fault detection will enhance the craft's ability to successfully

achieve mission goals. FPAs will consist of five major modules the main

controller, status monitor, fault diagnosis module, knowledge base, and

interface handler. These five modules will work together as an expert system

to detect errors early on and take preventative actions when possible. The

craft's autonomy will be increased by reducing the amount of ground control

intervention.

72
CCS is responsible for enabling all telemetry commands. It will also

process real-time commands and send them to AACS or FDS as needed,

monitoring them as well as other craft operations.

Section 5.2.2: ATTITUDE AND ARTICUL_,TION CONTROL SUBSYSTEM

¢

Utilizing an expert system such as APPS, AACS will hold the craft to its

chosen trajectory. In order to maintain communication with Earth, the high

gain antenna (HGA) must always point to Earth. This will be done by attitude

maintenance, antenna pointing, and gyro control. Antenna pointing control

will be divided into a main reflector and a subreflector drive.[13] More

information is contained in the Attitude and Articulation Subsection.

i

Section 5-2.3: FLIGHT DATA SUBSYSTEM

.

¢.

Responding to commands from the CCS, the Flight Data Subsystem

contains routines that control science instrumentation and the optical drive.

Some data processing is also handled through FDS.

FDS first collects engineering and science instrument data. This is used

to control the operation of the instruments. It is then formatted for either

storage or real-time transmission. Analog data must be converted to digital

form before it can be sent. This is also handled by the FDS. FDS must provide

data modes, rates and formats. FDS also provides frequency references for the

other subsystems.[ 19]

A system will be utilized using movable blocks of observations. These

groups are controlled relative to a single adjustable starting time, which

allows the computer to compensate for inability to determine the time of

closest approach in time for effective trajectory control maneuvers.[8]

Section 5-3: COMMUNICATIONS

Communications will consist of two subdivisions - Telemetry and the

Radio Frequency Subsystem (RFS). Ground commands will be processed

through RFS and passed on to CCS. Information is also relayed back to Earth

through the RFS by telemetry.

73
Section 5-3.1: TELEMETRY

Telemetry is made up of information from three different sources.

Science data is generated by instrument observations. While only small

amounts of science data are created, it requires the highest quality

transmission accuracy. Engineering data, for daily craft operation, requires a

moderate quality transmission of moderate volume. Imaging data, due to its

high redundancy, has a very high volume with the lowest quality standard.

In order to minimize data rates, all data that can be compressed shall be.

Both convolution and Reed-Solomon (RS) coding will be used, reducing bit

errors down to 10-6.[6] Although RS coding requires more processing at the

ground end, it is more effective and efficient than the Golay coding originally

used for the Voyager missions.[14] An RS system consists of one chip for the

encoder and seven chips for the decoder. It operates at a rate up to 80 Mbps.[3]

All, or nearly all pertinent data is retained.

Loss of information can also be reduced through multiple playbacks•

The memory capacity of the computers is high enough that the computer can

wait for data confirmation from Earth. If a high loss has occurred, data is

merely retransmitted.

Information will be sent in telemetry packets. Packets will be

constructed by individual subsystems. The CCS will add RS code bits to the

packets, providing error-free transmission.[6] Due to the large quantity of

memory, data loss can be further minimized.

74
Section 5-3.2: RADIO FREOUENCY SUBSYSTEM

Similar to the Cassini mission, Cerberus takes advantage of the more

powerful Ka-band (32 GHz). Although the antenna surface tolerances are

lower for higher frequencies, technology exists that will compensate for this.

This will include the addition of Ka feeds, waveguides and amplifiers. Ka-band

transmission simplifies hardware by decreasing size and power requirements.

Bit error rates and doppler tracking accuracy are also improved.J6]. A net gain

of 8 db can be realized by upgrading from X-band to Ka, as antenna gain

increases in proportion to the square of the link frequency. [l l] .

For further efficiency, the command detection unit (CDU) will be under

RFS jurisdiction. Its uplink command will be performed by the Ka-band

transponder. Technology exists allowing the telemetry modulation unit (TMU)

functions to be accomplished with a few chips. This too will come under the

RFS umbrella. [6]

Optical communications were also considered. Unfortunately, under the

proposal time limit, optical communication is not possible.[31] If developed, it

would have many advantages over the traditional RF system. Its wide

bandwidth would allow gigabits of information to be transmitted via a small

laser antenna with low transmitting power. However, the pointing accuracy

requirement alone, 10 -4 degrees, disqualifies it from use - the best accuracy

achieved today is 0.1 degrees. Output powers must be increased, as well as

improved beam quality. Materials must be found that are radiation tolerant.

Space debris is an additional factor, as it can damage the optics and will cause

an inferior signal quality.J4] For all these reasons, a traditional RF system was

judged to be best (Table 5-2).

Section 5-3.3: SPECIFICATIONS

For an RFS to meet the challenges of this mission, new techniques must

be employed. The Galileo antenna had the advantage of being made of a lightweight mesh material. Its folding capabilities allow many different launch

possibilities. However, in the ten years since the antenna was developed, new

technologies have provided different ways to improve upon its design.

Before we can use the Galileo antenna, it must be modified. New

technology allows us to transmit at the higher Ka-band frequency. Offset

75
i

i

r

;k .

subreflectors must also be added, to allow X- and S-band transmission. These

subreflectors will have frequency selective surfaces (FSS), allowing

transmission of several different frequencies from one antenna with the

addition of multiple feeds.[15,33] The antenna will be attached to a three-way

gimballed joint to allow it to be pointed in any direction. Sitting "on top" of the

craft, it will always be pointed at Earth, except in cases where the spacecraft

must be turned to perform science experimentation. The low gain antenna

(LGA) will also transmit at Ka-band. It will be used to communicate with Earth

until Jupiter is reached. The high gain antenna will be deployed at that time.

Table 5-2: Communication Options

Communication Mode

High gain antenna and Low

gain antenna

!Optical Communications

Disadvantage

Higher weight and cost.

Relatively high power

requirement.

Not yet developed.

Advantage

Higher redundancy,

allows delayed

deployment of HGA.

Proven technology.

Extremely lightweight

and power efficient.

Although bandwidths are traditionally 5-10% of the transmitting

frequency, the antenna is designed with a 21.85KHz bandwidth. This is

approximately 7E10"5% of the 32 GHz transmitting frequency. Additional feeds

must be used, as well as amplifiers and waveguides.

The antenna system must also be designed around a number of other

factors. DC to RF conversion (Lt or transmitter system losses) is currently

21%.[11] Attenuation, which increases with inclement weather, must also be

considered (Figure 5-1). Using Goldstone as the receiving station, one has an

atmospheric attenuation (La) of 92%.[27] Antenna efficiency, g,can be pushed

as high as 80%.[33] Other losses include receiver system losses (Lr), pointing

losses of both the transmitter and the receiver (Ltp and Lrp, respectively),

free space loss (Ls) and polarization loss between antennas (Lp). A

transmission power (Pt) of 10 W results in a receiving power of 9.05"10" 17

W.[30]

76
""1

"i

r= ,_ol

| l I l' | | I I I

BEST I

._ DSTONE

_AN_ERRA/MAORIO

oo_ I f 1 I I I I I

0 10 20 30 40 50 60 70 80 90

K= _AND _ 31SE TEMPE RATURE, K

f

i Fig.6-1: Ka-band atmospheric

noise temperature statistics: all

f sites, 30-deg elevation angle.[ 27]

fl__pOr:_OI;C

_._f_ec for

r , I/X ,_

Reflector

Fig.6-2: Multiple feeds for

different frequency bands can be

arranged with beam waveguides,

and frequency selective

surfaces. [ 15 ]

L,

DSN will be the receiving antenna. This requires a signal-to-noise ratio

(SNR) of at least 10, with higher values providing better transmission.[30] All

my calculations are based on the current 70m antenna with an SNR of 20. Vast

improvements in receiving power will be realized when the proposed array of

35m antennas is deployed. Other improvements could be made by orbiting a

receiving dish. Gravity effects would not be felt and atmospheric attenuation

would be eliminated. The dish could either be an orbiting satellite or a multideployable dish on the space station.

Section $-4: CONCLUSION

As few as ten years ago, the idea of a mission to Pluto would have been

ludicrous. It is only with recent advances that this mission has become

feasible. Without the ability to transmit at a higher frequency or use an

upgraded DSN, the antenna size alone would have precluded a successful

venture. Recent advances in materials such as Kevlar or optical disks further

decrease the size and weight of the craft. All of this adds up to a smaller, more

77
cost-efficient operation, able to explore not just Pluto, but other planets as

well.

78
References

¢

°

,

°

°

,

°

.

°

°

10.

11.

12.

13.

14.

15.

Adamski, Terrence,"Command and Control of the Voyager Spacecraft"

AIAA paper 87- 0501.

Casey W.L., Doughty G.R., "Laser Communications System Design," SPIE

vol. High Data Rate Atmospheric and Space Communications, Sept. 1988,

pp.28-35.

Cellarius, Mark R., "Data Compression for Data Acquisition,

Storage and PCM Transmission", Vol. 23, 1987 International

Telemetering Conference, San Diego Ca, pp. 517-24.

Chen C.C.,Schwartz J.A. "Space Station Based Deep-Space Optical

Communication Experiments", SPIE vol. 885 Free-Space Laser

Communication Technology, Jan. 1988, pp.2-9

Cox, P.A., "Knowledge-based Systems for the Hubble Space Telescope"

SPIE vol.1006, Space Station Automation, Nov. 1988, Cambridge Ma.

Draper, Ronald F. "The Mariner Mark II Program ", AIAA Paper

84-0214.

No.

Ford D.,Weeks D.,"Intelligent Data Reduction: Preliminary

Investigation"

Gray D.,"Navigating Neptune", I988 AIAA/AAS Astrodynamics

Conference Proceedings, Minneapolis Mn, pp.

Goforth, A., Dominy R., "AI, Automation and the Flight Telerobotic

Servicer", SPIE vol. 1006, Space Station Automation, Nov. 1988, Cambridge

Ma.

Hall, Justin R., McClure, Donald, "Voyager-Neptune Telemetry"

The DSN 70 Meter Antenna Upgrade", Vol. 23, 1987 International

Telemetering Conference, San Diego Ca, pp.163-70.

Hansen D.M., Kliore .J., ha-Band (32 Ghz) Benefits to Planned A • llV_

Missions" TDA Progress Report 42-88 N87-17958, 1988.

Kellogg, Kent, Spacecraft Antenna Group Supervisor, JPL, Pasadena Ca.,

private conversation, March 1990.

Kowakami Y, Hojo H., Ueba M., "Design of an Onboard Antenna

Pointing Control System for Communication Satellites", 1988

AIAA/AAS Astrodynamics Conference Proceedings, Minneapolis

Mn,pp. 689-694

Kosmann W.J., "Voyager-2 Neptune Encounter" Vol. 23, 1987

International Telemetering Conference, San Diego Ca, pp.117-23

Lee S.W., Houshmand B., Acosta B., "A Plate Mirror Antenna For

Earth Geostationary Platforms" NASA grant NAG 3-419

NASA

79
16.

17.

18.

19.

20.

21.

22.

23.

24.

25.

26.

27..

28.

29.

30.

Lee S.W.,Lo Y.T., The Antenna Handbook. Chapt.15,17,21,Van

Norston Press, New York,1988.

Lin, Luke, University of Illinois CS Grad Student, Private Interview,

April 1990.

Marderness, Howard P.,"Voyager Engineering Improvements for

Uranus Encounter" AAIA/AAS 1986 Astrodynamics Conference

Proceedines.

Madsen, Boyd D., "Voyager Neptune Telemetry System", Vol. 23, 1987

International Telemetering Conference, San Diego Ca, pp.127-36

Miller L.,Savary K., "Voyager Flight Engineering Preparations for

Neptune Encounter" 1988 AIAA/AAS Conference Proceedings,

Minneapolis MN, pp.351-361.

Miller, Warner, Marakis, James,"High Data Rate Reed-Solomon

Encoding and Decoding Using VSLI Technology", Vol. 23, 1987

International Telemetering Conference, San Diego Ca, pp.71-78.

Morris, Ray B., "Sequencing Voyager II for Uranus Encounter"1986

AIAA/AAS Astrodynamics Conference, Williamsburg Va.

Muratore, John F.,"Trends in Space Station Telemetry

Apllications", Vol. 23, 1987 International Telemetering Conference, San

Diego Ca, pp.ll-16.

Nicks, Oran W. F_r Travelers: The Exploring Machines Washington D.C.

1985

Pierce, John R., Posner, Edward C., Introduction to Communication

Science and Systems, Plenum Press, New York, 1980.

Ross M., ""History of Space Laser Communication", SPIE vol. 885 FreeSpace Laser Communication Technology, Jan. 1988, pp.2-9.

Slobin S.J.,"Models of Weather Effects on Noise Temperature and

Attenuation for Ka- and X-Band Telemetry Performance Analysis", TDA

Progress Report 42-88, N87-17961, 1986.

Smith J.G., "Ka-Band (32-GHz) Downlink Capability for Deep Space

Communications", TDA Progress Report 42-88, N87- 17956, 1986.

Sztipanouits J., Biegl C., Purves R.B., "Model-based Approach for

Intelligent ControI",SPIE vol.1006, Space Station Automation, Nov. 1988,

Cambridge Ma.

Wong, Carla M., Weeks, David J.,"Cooperating Expert Systems for

Station Power/Thermal Subsystem Testbeds" 1988 IECECConference

Proceedings, Denver Co.

Space

80
31.

32.

33.

conversation, March-April

Yuen, JosephH., Deep Space Telecommunications

Systems Engineering, Plenum Press, New York,1985

Yuen, Joseph H., JPL, Pasadena Ca., Private Conversation, April 1990.

Zimmerman, M., EM Graduate student University of Illinois,private

1990.

¢ -

J

Annendix E-1

f ,

eqtn. 5.1 Pr=l_ 16ARAT

PT Z,2L2

eqtn.5.2 G = 4nAT

t X2

eqtn.5.3 SNR=Pr/Pn

eqtn.5.4 Ls=(_/4nL) 2

eqtn.5.5

eqtn.5.6

eqtn.5.7

Pr Power received

Pt- Power transmitted

A R- Area of receiving

antenna

AT Area of transmitting

antenna

G T - Gain for transmitting

antenna

P - power of noise

n

SNR - signal to noise ratio

I.t - efficiency for antenna

L - distance between

spacecraft and ground

antenna

Pr=PtLtGTLTpLSLALpLRPGRLR

Pn=KTW

B =Wlog2(Pr/Pn+ 1 )

_/alue_;

SNR=20

LPluto=40A U* 149.6E9m

=5.984E12m

LRp

loss

LA - loss due to atmospheric

attenuation

LTP transmitter pointing

loss

receiver pointing

LR receiving system losses

LT transmitting system

losses

B - Bit rate of data

81
e

L Jupiter=5.203 A U * 149.6E 9m

=7.79E1 lm

_,=9.375E-3m

LR*GR*LRP=. 89

Lp=l.0

LA=.92

LTP=.89

LT=.21

Pt=10 W

T=3OK

transmission

LS- space loss

Lp - polarization loss

between antennas

G R - receiving antenna gain

K - Boltzmann's constant

T - temperature (in OK)

W Bandwidth (Hz)

Equation 5.1 gives a rough estimate of the antenna receiving power.

However, combining equations 5.2,.4 and .5 gives a more exact number:

HGA Pluto LGA at Jupiter

Pr=9.05E-17 W Pr--6.976E- 15

Using this and an SNR of 20, one can calculate bandwidth W from

equations 5.3 and 5.6. This, in turn, can be used to determine Pn:

WHGA = 9.05E- 17W/( (1.38E 10-23)30K(I 00) }

=21.85 kHz

Pn= 1.38E-23*3°K*21.85 kHz

= 9.05E-19 W

WLGA=1685 kHz

PnLGA = 6.976E-17 W

Using equation 6.7, one gets the information capacity:

B=21.85 kHz log2(100 +1)

=145.5 kbps

B=1685 kHz log2(100+l)

=11219.1 kbps

82
Section 6: SCIENCE SUBSYSTE M

!

F ,

Section 6-1: RFP REOUIREMENTS

The request for proposal (RFP) serves as a basis for the entire Cerberus

mission, yet the RFP includes only a few requirements that pertain to the

science subsystem. The most fundamental of these requirements is one which

states that the spacecraft must perform an unmanned study of Plutonian space.

This requirement defines the purpose of the entire mission and implies that

the science instruments aboard Cerberus must be suitable for studying Pluto,

its satellite, Charon, and the space surrounding the system. A related RFP

requirement demands that the spacecraft should be able to perform several

possible missions. This means that Cerberus' scientific instruments should not

be limited solely to a study of Pluto, but should also be useful for experiments

conducted elsewhere along the spacecraft's trajectory. For this reason,

Cerberus will gather data in interplanetary space, and if astronomers desire

more information about Jupiter, after completion of the Galileo mission, then

Cerberus will take measurements during its Jupiter flyby.

The RFP states that reliability, simplicity, and low cost must be

emphasized in the spacecraft design and mission planning. There are several

requirements which reflect this central objective. The first calls for

optimization of spacecraft performance, weight, and cost. For the science

subsystem, this requirement applies to the components selected for the

mission. In an effort to reduce costs and ensure reliability, the request for

proposal limits all components to off the shelf hardware available through

1999. To fulfill this requirement, Cerberus will primarily feature the

scientific instruments, or derivatives of these instruments, used on the

Voyager and Galileo missions. To further ensure mission reliability, each

instrument must have a sufficient design lifetime so that the instruments will

be functional throughout the mission and for a reasonable amount of time

afterwards. Since this lifetime will be on the order of twenty years, while the

Voyager mission is only thirteen years old and Galileo is just getting started,

the instruments have not yet been tested for the lifetime of the Cerberus

mission. These components, however, have undergone rigorous ground

testing and by the proposed launch date in 2002, they will have experienced at

least 12 years of flight testing, too. Finally, the RFP calls for artificial

83
intelligence to be used where applicable, providing the spacecraft with rapid

decision making capabilities while avoiding the long delays involved with

communication to and from Earth. Science applications for artificial

intelligence include automation of science instruments and data handling.

While the above requirements must be fulfilled by the Cerberus mission, there

is little restraint on the science which may be performed.

Section 6-21 METHOD OF ATTACK

Since the request for proposal sets so few standards for science

objectives, the selection of experimentation is at the discretion of the science

subsystem design engineer. In the past, studies of each planet have begun

with a flyby of that planet. For instance, Jupiter and Saturn were first studied,

up close, by flybys of Pioneers 10 and 11. Similarly, Uranus and Neptune were

first explored by Voyager II flybys. These missions serve as a model for the

experimentation to be carried out at Pluto. The information received from

Cerberus can then be used as a basis for further exploration of Pluto, just as

Galileo and Cassini will follow up where Pioneer and Voyager left off.

After it is determined which science will be performed, the selection

process begins for finding the equipment that will run the experiments. For

Cerberus, this procedure was accomplished by studying past, present, and

planned missions. The parameters taken into account included instrument

performance (spectral ranges, resolution capabilities, etc.), masses, power

consumption, and data rates. Table 6-1 lists the three latter parameters for

Cerberus' instruments. Amount of flight testing time was also considered in

the process. With the numerous variables to take into consideration, it was

difficult to perform numeric trade studies for each component, so the

instruments were chosen for their practicality and compatibility with a

mission of Cerberus' nature (inexpensive and reliable).

Insrument

SSI Camera

Table 6-1:

Mass

/kgt

29.7

Summar_¢ of Scientific Instruments

Data Rate (bps)

Power _W)

15.5 (ave.) 34.180 - 888,686

20.0 (max)

Temp.

Constraints

CCD -70deg C

(max)

84
?

Imaging

Spectrometer

Photopolarimeter/

Radiometer

Ultraviolent

Spectrometer

(18.0)

3.6

5.33

(8.0)

7.5 (Photo.)

4.5 (Radio.)

5.33

Magnetometer s 5.6 2.2 -

Plasma 9.9 8.1

Cosmic Ray

Radio Astronomy

/Plasma Wave

Radio Science

( ) Estimate

0

(lo)

6.7 (Radio)

1.1-1.6

(Plasma)

0

(500- 10,000)

180

266- 115,200

(Radio)

32 - 115,200

(Plasma)

0

No data available

Focal Plane

80 K (max)

-50 to +40

deg C

-18 to +6

deg C

-20 to +70

deg C

Ideally, it is desirable to perform as much science as possible on this

mission to Pluto. Due to limitations in power supply and instrument endurance,

though, the lengthy flight time will restrict the amount of data which can be

collected For this reason, the science sequence must be optimized by giving

priority to the most important experiments. Highest priority will go to the

remote sensing experiments at Pluto. Next in line are the particles and fields

studies at Pluto followed by the exploration of interplanetary space. Final

priority goes to the study of Jupiter. This type of method will help ensure that

the main objective of the mission, the exploration of Pluto, will be carried out

successfully.

Section 6-3: SCIENCE OBJECTIVES

Most of the information regarding the bodies of Pluto and Charon,

themselves, will be obtained from a series of remote sensing experiments. An

imaging device will take high resolution pictures from which studies will be

made providing information about the structure and motion of Pluto's

atmosphere (if it exists), as well as information regarding size, shape, color,

85
albedo, surface texture, and spin state. From these images, theories stating that

Pluto is covered with a methane haze can be tested and it can be determined

whether Pluto has a ring system. Spectroscopy studies will aid in the

determination of atmospheric and surface compositions. A photopolarimetry

experiment will reveal information about atmospheric particles and the

reflective properties of the surfaces of Pluto and Charon. Finally, a radiometer

will measure the visible and infrared radiation that is emitted and reflected by

the bodies so that the balance of energy between Pluto, its satellite, the sun,

and other sources may be studied.

Particles and fields experiments will probe the space around Pluto as

well as interplanetary space to enhance our knowledge of the solar wind and

its interaction with Pluto. Four magnetometers will measure magnetic field

intensity along the spacecraft trajectory allowing scientists to estimate the

shape and properties of Pluto's magnetic field. A plasma instrument will

identify and sample the energies and velocities of low energy ions and

electrons for studies of the solar wind and Pluto's magnetosphere. Properties

of cosmic rays will be tested by an instrument that measures the energies and

distribution of the high energy particles that make up these rays. The results

of this experiment should help scientists determine the origin and motion of

cosmic rays. Complimenting the plasma studies, a plasma wave experiment

will study the propagation of disturbances through plasma.

Other investigations include radio astronomy and radio science

experiments. Radio astronomy will involve the transmission to Earth of

various radio signals that Cerberus will encounter during its journey,

including those emitted by Pluto. The radio science experiment will utilize the

communication system aboard Cerberus in an effort to estimate the mass and

size of Pluto and Charon. The radio signals sent back from Cerberus, during

Earth occultation, will contribute to the investigation of atmosphereic density

and composition.

Section 6-4; COMPONENTS

With the science objectives determined, it is necessary to select the best

off the shelf hardware to run the experiments while conforming to mass, size,

cost, and power constraints. To perform the imaging, the solid state imaging

device (SSI) from Galileo will be inherited (See Fig. 6-1.). This system utilizes

86
an 800 x 800 element, charged-coupleddevice (CCD) with a silicon image sensor

array. Chosen for its high resolution capabilities, the SSI camera is more than

one hundred times as sensitive as the comparable vidicon-tube camera used on

Voyager.[1] This camera features an eight position filter wheel with filters

centered in the 727nm. and 889nm. methane absorption bands. It has four

exposure times and four repetition rates ranging from 2 1/3s. to 60 2/3s.[1]

The angular resolution of the SSI device is 8,128 mrad., and to minimize the

amount of data space necessaryfor each image, the camera has the ability to

compressdata at a ratio of about 2.5:1.[1] The CCD is sensitive to radiation and

operates at a maximum temperature of about -70 deg. C., so sensor shielding

and radiative cooling will be necessary.[1]

Visible and infrared spectroscopy measurementswill be taken by the

imaging spectrometer developed for the Mars Geoscience/

Climatology Orbiter (MGCO) (Fig. 6-2). This instrument was selected because, in

spite of the fact that it is scaled from Galileo's imaging spectrometer, it utilizes

superior sensor technology without additional weight or power requirements.

A silicon detector will sense visible and near infrared light ranging from 400

to 1000nm., and an indium antimonide sensor will be used for infrared

radiation

87
FIIELO

CO mmliCTO Pl

APE_qT_AIt El..! _*! NT$

CO IqME C-O Iq SI[CO_OAMY I

t _.l[ lilll %1?"S Ill I llll Pt0 IIII FIMIMARY FILTIM

IMI iq II10 Iq $1UlUT T E M

ELECTPlONIC$ CHAII$ i FOCAL lINT

Parameter

Angular resolution

Shortest exposure

Lon&est exposure

Active CCD area

Array aspect ratio

Pixel aspect ratio

Value

10.i6 ,r/pixel

4-t/6 ms

51.2 s

12.lg x [2.19 mm

I to i

1 to 1

Active lines per frame 800

Active pixels per line 800

CCD futt watt capacity I x I05 etectrons

Dark current <i0 electrons/s/pixel

B'its/picture element 8 raw

3.24 compressed

Readout noise _30 electrons rOOtmean-square/plxel

Number of filters 8

Gain states 4 (l. 4. 10. 40)

Mass 29.7 kg

Average power 15.5 W

Peak power 20.0 W

Volume L 90 cm

W 25 cm

H 30 cm

f

L Figure 6-1" SSI camera and parameters [1]

f

b ,

]

t

;-, --:..

, ---r...

Parameter

RITOI'Y-OIII'TII}I TELESCOPE

k

X

- RADIATIVE "_ _ .APERTUREl1 mr x 1 mr)

i ----_.... " _ -,' -" '_'I

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_ \_...._, t,_ OF POOR QUALITY

- k

x._. pL,_qETA,RY .---- AP_1_TUR£O00R

SHIELD

• "Value"

Design Altitude, km

Ground IFOV, m

Spectral Coverage, um

Spectral Sampling Interval, nm

Field of View, deg

Swath Width, km

Data Rate

Aperture Diameter, cm

Instruuuent T_aperature, K

: Signal-to-Noise Ratio at 3.5 um

: Signal-to-Noise Ratio at 2.0 um

, Focal Plane:

visible and Near Infrared (0.4 to 1.0 gm)

Short Wavelength Infrared (I.0 to ]°5 um)

Focal Plane TemEerature t K

300

300

0.4 to 3.5

20

7.7

40

Variable (Mode Dependent)

16.2

200

10.5

500

Silicon 32-Element Line Array

Indium Antimonide 128-Element Line Array

80

• Figure 6-2: MGCO imaging spectrometer and parameters [2]

• 88
between 1000 and 3500nm.[2] The MGCO spectrometer operates at a

temperature of 200K and the focal plane must be kept below 80K. so a

temperature control system must be implemented.J2]

A Photopolarimeter/Radiometer (PPR) device identical to that aboard

Galileo will be used on Cerberus (Fig. 6-3). This configuration is advantageous

because two experiments share the same equipment which reduces spacecraft

weight and complexity. The PPR must be kept at about 223K. and may not be

pointed toward the sun.[3]

The combination of Galileo's recent technology and flight experience

generally make its components the most attractive for use aboard Cerberus,

and like a majority of the other instruments, Cerberus' ultraviolet

spectrometer (UVS) will be inherited from Galileo. The UVS will examine light

ranging from l l5nm to 430nm.

1

This range exceeds 2 7 times that of Voyager's ultraviolet spectrometer.[4] This

instrument has two separate fields of view and covers its entire spectral range

in 4 1/3s.[4] This, and all of the remote sensing instruments, will be located,

together, on a platform of high pointing accuracy.

Four triaxial fluxgate magnetometers will be responsible for magnetic

field readings. These magnetometers were selected because their capabilities

have already been proven aboard Voyager.

The system will include two high field and low field magnetometers. Due to

their sensitivity, the low field magnetometers will be

suspended on a long boom, to avoid magnetic interference from the

spacecraft's RTGs and other components.

The superior energy range of Galileo's plasma instrument

(Fig. 6-4) makes it the best candidate for plasma studies aboard Cerberus. Its

range of 1.2 to 50,400V. is roughly ten times that of Pioneer's and Voyager's

instruments, plus its three mass spectrometers allow it to identify several

positive ions.[4] The instrument will be mounted on a rotating, low accuracy

platform so that it can analyze particle velocity distributions in all directions.

The energies of cosmic ray particles will be measured by the same type

of instrument as used on Voyager. The Voyager instrument was selected

because it was designed for a study of the

89
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90

ORIGINAL PP_GE IS

OF POOR QUALITY
outer planets and space beyond the solar system and heliosphere. The

extended energy range of 1-500MeV will allow the cosmic ray instrument to

study particles of higher energy that may exist in deep space.[5] The

instrument will accompany the plasma instrument on the low accuracy

platform.

Like Voyager, the design of Cerberus includes two long, perpendicular

antennas that will be used for both the plasma wave and astronomy

experiments. The electronics for the two experiments will be incorporated

into one component, thus conserving space and mass (Fig. 6-5). For the plasma

wave studies, the antennas will act as a dipole antenna, and for radio science,

they will serve as two monopoles.

The last experiment, radio science, requires no instrumentation.

Instead, it will use Cerberus' high gain antenna and communication system.

An analysis of perturbations in the spacecraft trajectory will yield mass

estimates for Pluto and Charon, while disturbances in the radio transmission to

Earth as Cerberus enters Earth occultation will reveal atmosphereic

properties. The length of occultation time will help astronomers better

estimate the size of Pluto.

Section 6-5: SCIENCE TIMEL|_E

The period of time dedicated to studying Pluto will last 50 days, centered

around the closest approach date. This period will begin with remote sensing

measurements taken at a low rate. At this time, the SSI camera will operate at

its slowest exposure time. About two weeks prior to closest approach, the

particles and fields instruments will begin to take readings at a much higher

rate than in interplanetary space as Cerberus approaches the magnetosphere

of the Pluto system. At about one week before closest approach, the frequency

of readings will have increased to a level which will necessitate a steady rate

of data transmission. The peak experimentation period will take place 24 hours

before and after closest approach. The camera will collect images at it highest

rate while operating at its fastest exposure time to avoid distortions due

91
J

|

i

t •

9'I
to the motion of the spacecraft. For a 12 to 18 hour interval during this busy

period, Cerberus' attitude will be adjusted so that the high accuracy platform

will point towards Charon and a detailed investigation will take place. The

precise timing for this maneuver will be determined when the configuration

of Pluto, Charon, Cerberus, and the sun, during encounter, is better known.

The positions of all of the previously mentioned bodies also effect the

sequencing of science events. For example, emitted radiation can only be

measured on the dark side of the body of interest. Earth occultation is another

subject which must be taken into consideration. When Cerberus enters

occultation, it will be impossible to transmit data to Earth, so all data will be

recorded for later transmission. A signal will still be sent by the spacecraft,

however, for radio science purposes. Following the 48 hour peak period, the

amount of data taking will be minimal, allowing time for the transmission of

recorded data.

The proposed flyby of Jupiter provides an excellent opportunity for

further experimentation. For a period of about 120 days, Cerberus will be in

the vicinity of Jupiter and could follow a science scheme similar to that at

Pluto. Galileo, however, is already on its way to an extensive exploration of

Jupiter, therefore, the results of the Galileo mission will dictate what science

will take place during the Jupiter flyby.

In interplanetary space, particles and fields readings will be taken at

regular intervals for solar wind studies. Prior to the Pluto encounter, data will

be collected every 1 A.U. Afterwards, the rate will be increased to once every

0.5 A.U. Using artificial intelligence, unknown bodies in Cerberus' path can

be explored as suspicious patterns in the particles and fields data will cause the

sensing instruments to search for large objects in deep space.

Section 6-6: INTERACTION WITH OTHER SUBSYSTEMS

The design of a science subsystem, for a complex project as Cerberus,

requires an extensive amount of communication between design team

members. The mission planning person must provide information on planet

and spacecraft dynamics, so that science sequencing can be arranged, and

submit details about space environment, so that the components may be

sufficiently protected. Since the mission revolves around the

experimentation, however, the science subsystem department is generally

93
/ -

responsible for providing information. The structures person requires the

masses of each component and their desired locations so that the spacecraft

design will maximize science performance. The science subsystem must work

with the AACS representative to ensure that instruments will be pointed with

the necessary amount accuracy. The CCC person must be informed about the

data taking scheme, including data rates, so that the communication system

will be able to process and transmit data with minimal losses. Automation of

science instruments is another concern shared by the science and

communications subsystems. Finally, the power and propulsion

representative must know about all of the component power requirements in

order to create a sufficient power system.

Section 6-7, FUTURE CONCERNS

This preliminary design sets up the basic concepts of the science

subsystem of Cerberus, but there are several details to be dealt with in the

following design phases. The extended flight time of the mission introduces

the question of design lifetime, yet by the proposed launch date in 2002,

Voyager's scientific instruments will have been operating in flight for 25

years and Galileo's instruments will have collected 12 years of experience.

Power and cost restrictions, as well as the constraints on data transmission,

will set a limit on the quantity of science that can be performed. These

limitations will have major effects on the details of the science timeline.

Another topic for the next design phase is the fulfillment of specific

instrument requirements. Heating, cooling, and shielding devices must be

implemented into to the spacecraft's design. These are just a few of the many

details which still need to be worked out. With the science objectives and

instruments intact, however, the preliminary design of Cerberus is a sound

one with no major "show-stoppers".

94
r-

[

i

[

t

r "

i

.

o

.

,

,

Numbered References

Klaasen, Kenneth P., Clary, Maurice C., and Janesick, James R.,

"Charged-Coupled Device Camera for the Galileo Jupiter

Orbiter Spacecraft", SPIE Vol. 493 The National Symposium

and Workshop on Optical Platforms. 1984.

Herring, M., Juergens, D. W., Kupferman, P. N., and Vane, G.,

"Imaging Spectrometer Concepts for Next-Generation Planetary

Missions", SPIE V91, 49_ The National Symposium and Workshop on

Qptical Platforms, 1984.

Watts, L. A., Russell, E. E., and Pellicori, S. F., "Photopolarimeter/

Radiometer Instrument for the Project Galileo Jupiter Orbiter",

SPIE Vol. 172 Instrumentation in Astronomy !II, Tuscon, Az.,

1979.

Yeates, C. M., Johnson, T. V., Colin, L., Fanale, F. P., Frank, L., and

Hunter, D. M., Galileo: Exploration of Jupiter's System, National

Aeronautics and Space Administration, Washington D. C., 1985.

Voyager N_ptune/Intcrst_llar Mission: Flight Science Office

$_ienc_ and Mission System_ Handbook. Jet Propulsion Laboratory, Oct.

1, 1987.

Other References

Report of the Comet Rendezvous Science Working Group. National

Aeronautics and Space Administration and Jet Propulsion Laboratory,

April 1, 1985.

Technology fgr Futur_ NASA Mi_i9n_; (_ivil Space T_chn91ogy

Initiative (CSTI) and Pathfinder, National Aeronautics and

Space Administration, 1988.

t

95