_91-16052

The Copernicus Project

Proposal for AAE 241

University of Illinios

April 24, 1990

Group Six:

Bob Barnstable

Hans Polte

Paul Kepes

Kevin Walker

Jeff Jacobs

Stephen Williams

https://ntrs.nasa.gov/search.jsp?R=19910006739 2019-05-09T14:36:09+00:00Z
i,!

r _:! L:

7_

.,.;,

Signatures

Bob Barnstable: Attitude and Articulation Control, Group Leader

Paul Kepes:

Hans Polte:

Mission Management, Planning and Costing

Science Instrumentation

5roB-o2 -/s 3

Jeff Jacobs: Power and Propulsion

Kevin Walker:

Stephen Williams:

Command, Control, and Communication

Structures

g

Z Zt,- Z7 -9 35ci
Executive Summary

The Copernicus spacecraft, to be launched on May 4, 2009, is

designed for scientific exploration of the planet Pluto. The main

objectives of this exploration is to accurately determine the mass,

density, and composition of the two bodies in the Pluto-Charon

system. A further goal of the exploration is to obtain precise images

of the system.

The spacecraft will be designed for three axis stability control. It

will use the latest technological advances to optimize the

performance, reliability, and cost of the spacecraft. Due to the long

duration of the mission, nominally 12.6 years, the spacecraft will be

powered by a long lasting radioactive power source. Although this

type of power may have some environmental drawbacks, currently it

is the only available source that is suitable for this mission.

The planned trajectory provides flybys of Jupiter and Saturn.

These flybys provide an opportunity for scientific study of these

planets in addition to Pluto. The information obtained on these

flybys will suppliment the data obtained by the Voyager and Galileo

missions.
Table of Contents

List of Acronyms

List of Tables

List of Figures

Introduction

Structures

Mission Management, Planning and Costing

Command, Control, and Communication

Power and Propulsion

Science Instrumentation

Attitude and Articulation Control

1-3

38-44

45-55

56-71

72-85

Conclusion 86 - 87
ACRONYMS

AACS: Attitude, Articulation, and Control System

ASTROS: Advanced Star/Target Reference Optical System

CCC: Command, Control, and Communication

CCD: Charge-Coupled Device

COM: Center of Mass

FDC: Flight Data Subsystem

FORS: Fiber Optic Rotation System

I-IECP: High Energy Charged Particle

HDA: Harmonic Drive Actuator

ISS: Imaging Science Subsystem

IIS: Infrared Interferometer Spectrometer

kbps: Kilobits Per Second

LECP: Low Energy Charged Particle

MITG: Modular Isotopic Thermoelectric Generator

MMPC: Mission Management, Planning and Costing

NAC: Narrow Angle Camera

PP: Plasma Panicle
ACRONYMS (cont.)

RCS: Reaction Control System

RTG: Radioisotope Thermoelectric Generator

TCM: Trajectory Correction Maneuver

S/C: Spacecraft

UVS: Ultraviolet Spectrometer

AV: Delta Velocity

WAC: Wide Angle Camera
List of Tables

Table:

1A Component Masses

1B Inertia of Copernicus

4A Power Requirements

5A Experimental Listing

5B Instrument Investigations

5C Instrument Data

6A Pointing Requirements

6B Scan Actuator Comparison

6C Star Tracker Comparison

6D Gyro Comparison

6E Thruster Location and Function Matrix

Page

lO

11

47

61

67

69

75

76

78

79

81
List of Figures

Figure:

1A

IB

IC Copernicus

2A Mission Timclinc

2B MULIMP Data

2C Copernicus Trajectory

4A Planetary Flyby Encounters

4B Power Breakdown

4C MITG

4D Thermoelectric Slice

4E Hydrazine Thruster

5A Power and Data Transmission Timeline

5B Instrument Use Timelinc

5C

6A Thruster Cluster Configuration

Page Number

6

6

7

23

25

26

48

50

51

51

53

60

60

68

81
INTRODUCTION
Introduction

The Copernicus Project proposal describes a Phase A design for an

unmanned mission to Plutoian space for the purpose of scientific

inquiry. This paper proposes that the spacecraft be designed, built,

and launched in an effort to increase our knowledge of the outer

Solar System and, in particular, the Pluto-Charon system. Thus far

Pluto is the only planet that has not been visited and investigated by

a space probe.

In order to insure an efficient and successful spacecraft and to

bring focus to the overall mission, the Copernicus Project proposal

will adhere to various mission guidelines and design requirements.

The following is a list of the spacecraft primary design requirements.

• The spacecraft must be unmanned.

• The spacecraft must be launched in the fh'st decade of the

twenty-first century.

• The spacecraft should be reliable and easy to operate.

• The spacecraft should use off the shelf hardware whenever

possible.

• The spacecraft should not use materials or techniques expected

to be available after 1999.

• On-orbit assembly should be identified and minimized.

• The launch vehicle to be used must be identified and the

interfaces must be compatible.

• The design must be flexible enough to perform several possible

missions.

• The design lifetime must be sufficient to carry out the mission

plus a reasonable safety margin.

• The spacecraft must use the latest advances in artificial

intelligence.

• The design will stress reliability, simplicity, and low cost.

• Four spacecraft will be built.

• Give an implementation plan for production of a final product.

2
In an effort to adhere to these design requirements and to create

an original and unique proposal, the project is divided into six

subsystems. Each subsystem is responsible for the design of a

specific area of the mission and the identification of any interactions

between the subsystems. An additional responsibility of each

subsystem is to optimize the performance, weight, and cost of the

individual subsystem in order to optimize those parameters for the

overall mission design. A list of the subsystems and their major

responsibilities follows.

Structures" Responsible for material selection for major spacecraft

components, component placement, thermal control for the

spacecraft, calculation of spacecraft inertia and center of mass, and

production planning.

Mi_iqn Mi_nagement. Planning and Costing: Responsible for mission

type selection, trajectory planning, launch vehicle selection, mission

timeline planning, and mission costing.

Cqmmi_nd. Control. and Communication" Responsible for the quality

of the spacecraft computers, the information storage capability of the

spacecraft, and insuring that the communication link with the

spacecraft is available at all times.

Power and Propulsion: Responsible for providing adequate power

supplies to the spacecraft components during all mission phases,

propellent selection, and propulsion unit selection and sizing.

$_i¢n¢¢ Instrumentation" Responsible for planning the mission

science objectives, planning the mission science timeline, and

scientific instrument selection.

Attitude and Articulation Control: Responsible for attitude control of

the spacecraft, maintaining antenna pointing requirements,

trajectory correction maneuvers, science maneuvers, and stability

throughout the mission.

3
STRUCTURES

4
Structure Subsystem: Introduction

The responsibility of the structure subsystem for the Pluto project

is to stress reliability, simplicity, and low cost in the areas of material

selection, thermal control, and overall spacecraft design. Subjects to

consider in fulfilling this responsibility are minimizing the spacecraft

weight, minimizing the amount of on-orbit assembly of the

spacecraft, and insuring a design lifetime sufficient to carry out the

mission plus a safety margin. An additional responsibility is to

provide an implementation plan for production of the final product.

To meet these requirements the structure subsystem is divided into

the following areas of consideration:

1. Drawings of the spacecraft

2. Placement of the spacecraft components to

meet requirements

3. Mass and inertia of the spacecraft

4. Material selection

5. Thermal control

6. Launch vehicle compatibility

7. On-orbit assembly

8. Production of the final product

9. Interactions with other subsystems

Drawings of the Spacecraft

Drawings of the spacecraft are provided to enhance the reader's

conception of the component placement and the overall spacecraft

design. The major spacecraft components included in the drawings

are the bus, propellent tank, main propulsive unit, three boom

extensions, RTG, scan platform, and antenna unit. Major spacecraft

dimensions are provided in meters. Two views of the spacecraft will

provide the reader with a clear idea of the spacecraft configuration.

The spacecraft axis was selected such that the origin coincides

with the geometric center of the bus. The Z-axis points out along the

antenna mast, the X-axis points out along the magnetometer boom,

and the Y-axis points out along the science boom to form a standard

righthanded coordinate system.

5
RTG /

I

3m

10 °

Magnetometer /

boom

Figure 1A

13m

J

Scan

platform

I I

I I l

I l

I I

antenna

[ !

Bus

Figure 1B

X

Propellant

tank

D

6
Figure IC

Copernicus

7
LJ

Placement of Spacecraft Components to Meet

Requirements

The driving requirements of this mission are reliability,

simplicity, and low cost. From these primary requirements come

several derived requirements that influence the positioning of the

individual spacecraft components. These derived requirements are:

radiation protection for all spacecraft components, the scientific

instruments must have a clear field of view, no component that

would disrupt communications should be placed near the antenna,

the main propulsive unit should create a line of force through the

spacecraft center of mass (COM), the components in the bus must be

compact to aid in thermal control of the bus, and the magnetometer

must be isolated from the interference of other spacecraft

components.

The radioisotope thermoelectric generator (RTG) emits radiation

that is damaging to other spacecraft components. To minimize this

radiation damage the RTG should be placed as far as possible from all

other spacecraft components. The distance that the RTG can be

placed from the main spacecraft structure is limited by the strength

of the RTG boom and spacecraft COM considerations. In an effort to

keep the COM on the Z-axis for attitude control simplification, the

RTG will be placed approximately 3 meters from the bus and at an

angle of 10 ° off the negative Y-axis in the XY plane. For additional

radiation protection, a metal shield will be placed at the end of the

RTG boom between the RTG and the main body of the spacecraft.

The placement of the scan platform must provide an adequate

viewing range for the scientific instruments. This is one of the most

important placement requirements. If this requirement is not met,

then the success of the mission will be limited. The scan platform

will be placed on a 1.2m boom that extends 0.3m beyond the rim of

the antenna. This placement was achieved by a tradeoff of field of

view and the previously mentioned COM restriction. Also, the scan

platform will be placed such that the spacecraft main body is

8
between the platform and the RTG for redundant radiation

protection.

Communication is also essential for the success of the mission. In

an effort to increase reliability, any components that are placed

within the antenna's field of transmission or reception should be

transparent to the antenna. A better placement technique is to leave

this area of the antenna free of any components at all. The second

technique is simpler than the use of antenna transparent components

and it was therefore selected.

To prevent any unwanted torques while the main propulsive unit

is in operation, the unit will be oriented so that its line of force

coincides with the Z-axis of the spacecraft. As previously stated, all

spacecraft components will be positioned so that the spacecraft COM

lies on the Z-axis.

The components housed within the bus will be placed in a

compact manner. This technique reduces the overall volume of the

bus and therefore the volume that requires the most thermal control.

The method in which this reduction in thermal control cost is

achieved will be discussed in a later section. The compact placement

of the components within the bus helps to reduce the mission cost

and thereby helps to fulfill a primary mission requirement.

A final placement requirement involves the magnetometer. The

magnetometer must be placed as far as possible from the other

spacecraft components to reduce the amount of interference

encountered from the other components. Again, the distance that the

magnetometer can be placed from the main spacecraft assembly is

restricted by COM placement, the strength of the magnetometer

boom, and the cost per unit length of the boom.

Mass and Inertia of the Spacecraft

Mass estimates are provided only for the major components of

the spacecraft. The following mass estimates are derived from other

subsystem requirements, considerations, and calculations.

9
Table 1.A. Component Masses

Component: Mass (kg):

Antenna 5

Antenna Base 45

Bus (includes Structure,

Thermal Control, and Cabling) 270

Computers 100

Science Platform 111

Science Boom 35

Magnetometer Boom 5

RTG Boom 5

RTG 60

Propulsion Unit Tank 120

Propellent 1500- 2000

Total spacecraft mass (unfuelled): 756 kg

The inertia of the spacecraft is calculated with the aid of a

computer program. The inertia and COM of individual components

are calculated by hand and these results are input into the program

which calculates the overall spacecraft inertia and COM. The

individual components are idealized into geometric shapes to

simplify the inertia calculations as described in the structure section

appendix.

In an effort to simplify the placement of the attitude thrusters

and the main propulsive unit, the spacecraft COM should lie on the Zaxis and as close to the geometric center of the bus as possible.

Several trials were performed in which the lengths of the science

boom and the RTG boom were varied. An additional variable was the

angle between the RTG boom and the negative Y-axis in the XY plane.

On the ninth trial the spacecraft COM was within approximately 0.5

cm of the Z-axis and approximately 11 cm below the geometric

center of the bus. This result was obtained with the unfuelled

configuration. This position of the COM is adequate for the purposes

of this preliminary design report.

10
The inertia and COM for the unfuelled configuration of the ninth

trial is:

Table 1-B.

Body Name: Copernicus

Inertia Matrix: 2334.0560

-155.8637

-.3384

Inertia of Copernicus

-155.8637 -.3384

700.0375 -.4187

-.4187 2724.9290

Body COM: .0039

Body Mass: 756.0000

Number of Bodies: 9

.0048 -. 1147

Princip_InertiaMatrix:

2348.7910 .0000 .0000

.0000 685.3029 .0000

.0000 .0000 2724.9290

Eigenvector Matrix:

.9956 .0941 - .0008

-.0941 .9956 - .0001

.0008 .0002 1.0000

Material Selection

There are several factors to consider in the material selection

process. First, to comply with the primary mission requirements, the

materials should be light weight, low cost, and should reliably fulfill

their design function. Additional material selection considerations

include: radiation damage threshold, contamination resistance,

thermal characteristics, strength, stiffness, and general structural

qualities. These characteristics must be carefully considered when

selecting materials for the spacecraft.

The main purpose of this mission is scientific exploration of

Plutoian space. Therefore it is essential that the science instruments

11
be kept operational. Contaminants such as atomic oxygen, outgassed

materials, and cosmic debris will accumulate on instrument surfaces

over time and impede their performance. Since the mission is of

such long duration, contamination protection of the instruments must

be a major factor in material selection.

One method of protecting the instruments is by installing a

permanent cover which is transparent to the instrument. A second

means of protection is the retractable cover design. This design

involves moving parts and should be used only where absolutely

necessary in an effort to enhance simplicity. If the retractable cover

should fail to open, then the success of the mission would be limited.

A redundant method of radiation protection is achieved by

placing a metal shield between the RTG and the main spacecraft

body. An aluminum shield was selected due to its low cost, light

weight, and high radiation damage threshold. Composites should not

be used for this application due to their susceptibility to radiation

damage. 1

An application that is well suited for composite materials is the

main antenna. The composite can be easily molded into the unique

antenna shape. Also, because of their low coefficient of thermal

expansion and high thermal conductivity, composites can be used in

systems which require high thermostructural stability like the

antenna dish. 2

For the main structural supports of the spacecraft, titanium

should be used where strength and thermal stability is important.

Graphite epoxy can be used in secondary truss supports and

stiffeners. Aluminum is attractive for its strength to weight ratio,

availability, low cost, and because it is space proven.

In situations where the stiffness of a structural member is crucial,

beryllium will be used instead of titanium. The modulus of elasticity

of beryllium is 2.5 times that of titanium and beryllium is

considerably lighter in weight. Although beryllium is more costly to

produce than titanium, beryllium's weight savings makes it less

costly than titanium to put into orbit. 3

The use of cosmic ray resistant parts for the computer's electronic

components will depend on their performance on the Galileo probe.

12
Sandia National Laboratories developed these components in an

effort to reduce the number of single event upsets in the computer's

logic and memory. 4 If these components prove successful in

reducing the number of computer sequence failures and if the cost is

reasonable, then cosmic ray resistant parts should be incorporated

into the Pluto probe's computer for enhanced reliability and

performance.

Thermal Control

Thermal control will insure that each part of the spacecraft will

have an appropriate thermal environment for operation. The

different components will require significantly different thermal

environments so that temperature gradients will be present

throughout the spacecraft. Thermal control will be further

complicated by the changing thermal surroundings as the mission

progresses. The three most significant phases are: thermal control

on Earth and during launch, thermal control in space close to the sun

(0.5-3 AU), and thermal control in the outer solar system.

The problem of thermal control is best solved by examining the

major components of the spacecraft.

Bus: The major considerations for thermal control of the bus are

isolation from solar heating, internal coupling to prevent

temperature gradients, and heat rejection at external bus surfaces. 5

A very cost and weight efficient method of preventing solar heating

in the bus is by the use of multilayer insulation blankets. This

passive thermal control technique makes use of the unique insulation

properties of multilayer designs. Redundancy is also achieved by

using multiple layers. The material is selected for minimum heat

transmission except for a few layers of very tough material such as

Teflon for micrometeoroid protection. The internal coupling is

achieved by positioning the internal components as compactly as

possible. This technique produces a smaller volume to be thermally

controlled and thus the cost of thermal control is reduced. This helps

meet the low cost mission requirement. The heat rejection phase is

13
ri •

,._.j

U

U

accomplished by transporting waste heat from the interior of the bus

to the external bus surfaces via a system of thermal switches. At

points along the external bus surface are heat radiators in the form

of thermostatically controlled louvers. There will be several of these

louver sites for redundancy.

RTG: The RTG produces large amounts of heat to be converted

into electrical power for the spacecraft. Due to radiation protection

considerations, the RTG is relatively isolated from all other

spacecraft components. This isolation also serves as an excellent

thermal barrier between the RTG and the spacecraft. Any waste heat

produced by the RTG can easily be rejected into space by an array of

metal fins that act as passive heat radiators.

_: The hydrazine thrusters will be thermally controlled

by strip heaters constructed of printed heating element circuits

imbedded in Kapton film. 6 These heaters will be placed on the

catalyst bed of the thrusters to produce temperatures well above

500K. The hydrazine fuel lines will be heated by wrapping wire

heating elements around the fuel line.

Science Instruments: The great design flexibility of the printed

circuit strip heaters mentioned above will allow them to provide

thermal control to the science instruments as well as the thrusters.

The design temperature for the science instruments is approximately

140K which is well within the thermal range of the strip heaters. To

help meet the requirement of redundancy in all spacecraft systems,

two strip heaters will be provided for every science instrument and

every thruster. This increase in thermal control should not produce

a drastic increase in overall spacecraft weight due to the very small

mass of the strip heaters.

Launch Vehicle Compatibility

The spacecraft must be compatible with the selected launch

vehicle. This means that all interfaces between the spacecraft

and the launch vehicle must be selected for compatibility. Also,

14
the dimensions of the spacecraft cannot exceed the payload

dimensions of the chosen launch vehicle.

The launch configuration is approximately cylindrical in shape.

The approximate dimensions of this cylinder are: width-3.7 m and

length - 4.5 m. The width r- ponds to the antenna diameter.

The antenna is

antenna used on

is not folda bl

The leng

payload az _

The

create any

thermal conu

control while

the launch ve

include power

mountings that h

vehicle. All these

a one pip- and is similar in design to the

"t. The fact that the antenna

inch vehicle selection at all.

-,ate any problems with the

launch vehicle must not

e interfaces include

assistance in thermal

rich pad and while

3ther interfaces may

interfaces, and the

position inside the launch

oe compatible.

On-Orbit Assembly

ORIGINAL PAGE F3

OF POOR QUALITY

The Copernicus will be a complete unit in its launch configuration.

No assembly will be required while in orbit. However, there will be

several boom deployments and general transformations of the

spacecraft from its launch configuration to its cruise configuration

while in orbit. Separation from the launch vehicle and upper stage

will be achieved by pyrotechnic methods such as explosive bolts.

These deployments will be made while the spacecraft is in LEO.

This will enable a repair and/or rescue attempt in the event of a

deployment failure. If the deployments are made in GEO or on route

to Pluto and a deployment failure occurs, then repair attempts would

be much more difficult to engineer. Deployment of the booms in LEO

will help improve mission reliability.

15
Production of the Final Product

The production of the Pluto probe will be a multistep process of

design, parts construction, system integration, and possible redesign.

In each of these phases testing for quality and reliability is essential.

A series of testing procedures has been described that helps insure

the production of reliable spacecraft. 7 The following is a description

of that testing procedure.

Test Objectives:

Development Test: Establish a fundamental behavior pattern upon which a

design

meet all

can be based.

Qualification Test: Verify that the equipment and associated software will

, specified requirements.

Acceptance Test: Verify worknmnship and demonstrate that the

equipment

functions properly over the range of correctly selected operating conditions.

Prelaunch Verification Test: Performed at the launch site to verify that

the

mated to

spacecraft has sustained no shipping damage and has been properly

the Munch vehicle.

Interactions with other Subsystems

Mission Planning: The dimensions of the spacecraft in launch

configuration limits the mission planner's selection of launch vehicle.

Also, the mission planner has selected a flyby mission which greatly

simplifies the overall spacecraft configuration.

16
Science: The scanning and pointing requirements of the scientific

instruments requires that the scan platform be positioned in a clear

field of view. The scientific instruments must also b¢ provided with

shielding from the contaminating space environment. Thermal

control must be provided.

Attitude and Articulation Control: The spacecraft inertia and COM,

determined by component masses and positions, affects the

placement of attitude thrusters and thruster force selection. Thermal

control must be provided.

Command, Control. and Communication: The antenna size and

placement places restrictions on the placement of the scan platform

for clear viewing. The massive computers housed in the bus

significantly affect the spacecraft inertia and COM. Also, the

computers generate heat that must be rejected from the bus by

radiating louvers.

power and Propulsion: The propellent tank, when fuelled, is the

most significant factor in determining the spacecraft inertia and COM.

The main propulsive unit must be oriented so that its line of force

acts through the spacecraft COM. Thermal control must be provided.

17
Appendix 1A: Inertia Calculations

The calculation of the individual component inertia's is simplified

greatly by idealizing those components into simple geometric shapes.

This assumption yields results which are adequate for the purposes

of this preliminary design report. Of course this simplified

methodology is in no way appropriate for actual inertia calculations

of the later stages of design. Another simplifying assumption is that

each component is homogeneous in density. Also, for the purpose of

this calculation the mass of the bus includes the bus structure,

command and control computers, thermal control, and cabling. The

following is a list of component idealizations, component inertias, and

component COMs. All dimensions are in meters. All inertias are in

units of kg-m 2.

Bus (370 kg): Hollow cylinder. Lffi.35 Roffi.95 Riffi.65

IxffiIyffiM[(Ro2+Ri2)/4+(L2)/12]-- 126.3

Iz-M(Ro2+Ri2)/'2ffi245

COMffi(O,O,O)

Propellent Tank (120 kg empty): Spherical shell.

Ix=IyffiIzffi2MR2/3=80

R=I.0 COMffi(0,0,-.94)

Antenna (5 kg): Flat disk. R=1.85

Ix=Iy=MR2/4=4.3

Iz=MR2/2=8.7

COM-(O,O,.7)

Antenna Base (45 kg): Solid cylinder. Lffi.12

Lx=Iy=M[(R2)/4+(L2)/12]ffi 10.2

Iz=MR2/'2=20.3

R=.95 COM--(0,0,.5)

Magnetometer Boom (5 kg): Thin rod. L=13 COM=(7.45,0,0)

18
Ix={)

Iy=lz=ML2/12=70.4

RTG Boom (5 kg): Thin Rod

Iy=0

Ix=Iz=ML2/12=3.7

L=3

Science Boom (35 kg): Thin rod.

iy=0

Ix=Iz=ML2112=4.2

L=l.2

Scan Platform (111 kg): Prism. L=.5

Ix=M(W2+I-I2)/12=I.7

Iy=Iz=M(W2+L2)/12=3.1

W=.3

RTG (60 kg): Cylinder. R-.1 L=1.52

IyfMR2/2=.3

Ix=Iz=M[(R2)/4+(L2)/I2]= I1.7

H'-.3

COM--(-.5,-2.45,0)

COM=(0, 1.55,0)

COM=(0,2.2,0)

COM-(-.53,-4.71,0)

19
.

.

.

1

t

.

.

Appendix 1B: References

"Configuring the Grand Tour Spacecraft." Advances in the

Aeronautical Sciences, vol. 29, part 2. p. 447.

"Structures and Materials Technology." NASA

Technoloev Model. June, 1985. p. 2-12.

Space Systems

"Beryllium." Machine Design, June 7, 1984. p.59.

"Galileo to Employ Cosmic Ray-Resistant Parts." Aviation Week

and Space Technology, May 7, 1984. p. 20.

Dumas, L. N. "Temperature Control of the Mariner Class

Spacecraft: A Seven Mission Summary."

A_[r0nautics and Aeronautics, vol. 35: Thermophysics and

Spacecraft Thermal Control, 1974. p. 538.

Taylor, Charles H. "Precise Touches of Warmth in Cold Space."

_America, October, 1984. p. 28.

Krausz, A. "The Effectiveness of Spacecraft Testing Programs."

Journal of Environmental Science, January, 1980. p. 14-18.

20
MISSION MANAGEMENT,

PLANNING,

AND

COSTING

21
5

D

Introduction

Mission Management, Planning and Costing (MMPC) has several

responsibilities regarding the unmanned mission to Pluto. A mission

timeline, outlining such features as the launch date, impulse points,

encounters with planets, arrival at Pluto, and the proposed end of

mission date must be furnished. MMPC must also determine a

trajectory system so that time and A V are optimized. Another

responsibility is the selection of the launch vehicle. A vehicle which

minimally satisfies the spacecraft's dimensions at launch as well as

the mass of the launch configuration is necessary. Furthermore,

MMPC must also select the type of mission to be performed at Pluto.

The mission should stress simplicity, reliablity and low cost. Lastly, a

total costing analysis for the project must be furnished.

The remainder of the MMPC section contains a detailed analysis of

the requirements previously mentioned, including trade studies and

mission planning effects on other subsystems. The requirements are

treated as separate categories where applicable, and each will be

discussed individually.

Mission Timeline

On May 4, 2009 (day 0) NASA will launch the spacecraft

Copernicus into a low earth orbit (LEO) of 270 km and an eccentricity

of 0.00. The spacecraft will then leave the Earth's orbit via an upper

stage and begin it's voyage to Pluto. On March 1, 2010 (day 300.4)

Copernicus will fire an impulse to prepare for its gravity assist at

Jupiter. This gravity assist at Jupiter will occur on February 18,

2012 (day 1019.8). The spacecraft will then be on a trajectory for

the planet Saturn, arriving on July 29, 2015 (day 2276.6). Once

again, a gravity assist will be made. Copernicus will then travel

uninterrupted for about six years until it reaches its target

22
destination, Pluto. The spacecraft will fly by Pluto on December 14,

2021 (day 4607.0). It will then continue on, leaving our solar

system, not to return. The end of the mission will occur after the

encounter with Pluto on December 14, 2021 (day 4607.0).

During it's flight, Copernicus will be performing correction

maneuvers (see Attitude and Articulation Control) when necessary.

As they cannot be predicted, no mention of it is included in this time

schedule. Figure 2.A shows a timeline view of the mission, from the

launch date to the end of mission date.

MMPC timeline

impulse Saturn flyby

2010

launch Jupiter flyby

May 4, 2009 February 18, 2012

Pluto arrival

December 14, 2021

Figure_.A. Mission Timeline

The overall duration of the mission is 12.613 years (4607.0 days).

During this time a management program will be in effect. The

structure of this program will include a management, control,

administration and support staff as well as division representatives 1.

Also, the duration time pertains only to flight of the spacecraft and

does not include the planning, research and development and the

assembly and testing.

23
Trajectory Systems

The selection of a trajectory system is perhaps the biggest task

for the MMPC subsystem. The spacecraft ideally should arrive at

Pluto in a minimum amount of time, while using a minimum amount

of fuel. This immediately produces a conflict. A compromise which

effectively minimizes both is desired.

The analysis of a trajectory system was performed with computer

software. The spacecraft had the requirement that it must be

launched sometime in the first decade of the twenty-first century.

The spacecraft would have to travel about 33 AU's. A direct flight to

Pluto was on the order of 28 years 2. This was double the desired

flight time so efforts to use gravity assists were employed. The first

system consisted of using Jupiter as a gravity assist. Much work cut

the flight time down considerably to about 15-16 years 2. However,

more planet gravity assists to shorten the flight time were still

necessary. The next project involved using Jupiter, Saturn and

Neptune for gravity assists. The project was named EJSNP (EarthJupiter-Saturn-Neptune-Pluto). The project was aborted in one

week. Neptune could not line up properly in conjunction with the

other planets, and was requiring too large a A V to correct it. So

project Pluto began, consisting of an "Earth-Jupiter-Saturn-Pluto"

configuration. This cut the flight time down on the order of 13-13.5

years. However, one problem was that it is desirable to leave before

or after the first decade of the twenty-first century for Saturn and

Jupiter to align properly, preferably early. Another problem is that

Jupiter is not to be approached closer than 10 body radii due to large

radiation output and Saturn should not be approached closer than 2.4

body radii due to it's rings. The Jupiter restriction was not a problem

but Saturn continually required an approach of less than two body

radii. The project was switched to "Longshot", using the same bodies

as project Pluto but using a launch time at the end of the decade.

This allowed Saturn's restriction to be satisfied and produced a flight

time of about 12.8-13.2 years. The following graph (Figure 2B)

depicts a trade off between A V required and time for Operation

24
Longshot.

2C.).

The final trajectory selection was then determined (Figure

g.

2g

|

I

0

"@

Data from "Mulimp"

22

20

18

16

14

12

10

11 12 13

time (years)

4

Figure 2B. Mulimp Data

The final selection was optimized to produce an impulse to assist

the gravity assist at Jupiter. The mission flight time was finally

reduced to 12.613 years. The complete analysis of this mission can

be found in the appendix after this subsystem, including but not

limited to launch time, A V required, and the coordinates of the

specific events. The final trajectory is mapped in Figure 2C. Note

that planet sizes are not to scale but are shown for illustration

purposes.

Another problem is the solar system's asteroid belt. To avoid any

possible collision that might result in a mission failure, the impulse

fired after departure will provide a A V of 0.267 km/sec in the

negative z-direction (see Appendix). Another advantage with this

trajectory is that it uses all of its fuel (not including the safety factor

25
Sun

Earth

Jupiter

Impulse point

Pluto

Figure 2C. Copernicus Trajectory

26
of fuel) early. After 2.798 years, Copernicus will have made its last

burn and will travel the remainder of its ten years with the weight

of the spacecraft only (excluding attitude control fuel). This is also

responsible for its short flight time. The total A V required for the

mission is 12.371 km/sec. This includes the departure from the

Earth's orbit. The A V required from the spacecraft's propulsion

module is 6.123 km/sec. This can also be found in the Appendix

following this subsystem.

Launch vehicle selection

The spacecraft Copernicus requires a launch vehicle to insert it

into earth's orbit. The selection of vehicles was limited to United

States launch vehicles. The launch vehicle would have to be able not

only to reach orbit, but it was also desired to use a configuration that

would let the spacecraft escape the earth's gravity and to begin it's

mission.

The launch vehicle must satisfy the spacecraft's weight including

fuel and launch packing. A factor of safety of at least l0 percent was

also desired. The companies that were considered were Martin

Marietta, General Dynamics (GD)/Space Systems, McDonnell Douglas,

and Boeing. Initially, GD/Space System's Atlas G was selected.

However, as more fuel was added, the spacecraft's weight increased

and the minimum performance payload necessary became 2730 kg

and the Atlas G could no longer meet the requirement 4. The

spacecraft's pre-launch configuration had the dimensions of a

cylinder of radius 3.7 m and a height of 5.0 m (see "Structures").

These dimensions could be employed on most launch vehicles and

was not a primary concern.

The launch vehicle finally selected consisted of the vehicle and an

upper stage. The launch vehicle selected is the Titan T-34D, by

Martin Marietta. This, used in combination with the Centaur D1-T,

could handle a payload weighing up to 5910 kg 4. This exceeds the

minimum requirement easily. However, the Centaur upper stage is

27
built by GD/Space Systems and the Centaur D1-T is a modified

version of the Centaur, designed specifically for the Titan T-34D.

The Titan T-34D uses a solid propellant while the Centaur D1-T

uses LOX/LHz Both vehicles are environmentally safe and pose no

conflicts regarding the safety of the launch.

Mission Type

The type of mission selected is the result of a lengthy trade

analysis. The types of missions were divided into three categories:

flyby, lander and orbiter. A flyby class mission was identified as any

mission which did not perform any thrusting at Pluto. A lander

mission was defined as the landing of any item on Pluto's surface.

Lastly, an orbiter mission involved using a burn to obtain an orbit

about the planet for a given length of time. Of the three classes, only

the flyby and the orbiter missions were highly analyzed.

A lander mission involved sending a spacecraft to a planet of

which there is little knowledge of. Historically, a lander mission

follows an initial study of the planet. For a lander mission to be

effective, an accurate idea of what is to be accomplished should be

known. It would be senseless to send a lander to Pluto without first

knowing what areas of the planet interest us. Also, the difficulties of

uncertain areas including the gravity, composition, surface conditions

and temperatures possess too high a risk factor for such a mission.

Furthermore, the cost of carrying out a lander mission to Pluto might

as much as double that of a flyby.

Initially, ideas for an orbiter mission were assembled. An orbiter

could perform many experiments, and would also allow a longer

encounter time at Pluto. Also, the mission was to incorporate a

needle probe to penetrate the surface of Pluto and to examine

samples. However, the AV required was high (a hum of 9.0-11.5

km/sec was required to insert the spacecraft into orbit2). Also,

further research posed yet a bigger problem: Pluto's moon, Charon.

This was of no major concern at first. However, since Charon's

mass exceeds 4 percent of Pluto's mass 6, the two bodies behave as a

28
binary system. This system would make an attempt to orbit Pluto

very difficult. Essentially, a three body problem must be solved.

Another idea would be to orbit in a "figure eight" configuration. Also,

while Charon's sphere of influence is estimated at 7000 km , Pluto

can retain a satellite up to an estimated 5000000 kin6. The mission

tended to lean toward flyby class at this point. To orbit Pluto, burns

would likely be needed for stable equilibrium. This suggests that the

orbit duration would be short (finite), and a finite orbit duration did

not warrant the increased cost of fuel required for orbit insertion.

The mission to send an orbiter to Pluto was finally aborted. A

mission to flyby Pluto was decided.

A flyby mission is the least expensive to build, test and fly. The

components needed for the mission are considerably less than that of

an orbiter class mission, making it a simpler design and more

reliable. Furthermore, a flyby mission has an attractive A V (see

"trajectory system"). While a flyby mission has less of an

opportunity to gather information, it still provided adequate

instrumentation, including imaging equipment to make an initial

survey of the planet. Lastly, the spacecraft would ideally leave the

solar system permanently. The spacecraft will have drawings on it's

buss including a picture of man, as well as the location in our solar

system in the Milky Way galaxy in the event of an encounter with

any intelligent life. Only a flyby mission would allow this to occur.

Costing

The costing of the spacecraft includes the cost of not only the

design and research leading to the construction of the vehicle, but

the ground support operations of the lifetime of the mission. A

detailed analysis of the costing can be found in the appendix

following the end of this subsystem. The costing estimation used in

this report is the "model estimation" method. This primarily involves

assigning a number of labor hours to each section of the spacecraft.

The labor hours are in turn converted into labor cost and the labor

29
cost is finally related to the total cost. The total cost of the spacecraft

is $999,443,600 dollars in terms of the 1977 fiscal year.

Another estimation technique is the concept of inheritance. The

model estimation technique uses the masses of the individual

systems but gives no consideration to the design and research

development of the systems. Inheritance involves assigning each

system to one of five classes:

Class One:

Class Two:

Class Three:

Class Four:

Class Five:

Off-the-Shelf/Block Buy

Exact Repeat of Subsystem

Minor Modifications of Subsystem

Major Modifications of Subsystem

New Subsystem

Any components from class one will benefit from the previous

design while class five receives no benefits whatsoever. By

incorporating inheritance into the model estimation technique, the

final cost will be effectively estimated. Assume that four spacecraft

will be built for costing purposes.

30
Appendix 2A.

CostingSection 1:

This contains the determination of direct labor hours (DLH) and

recurring labor hours (RLH). The standard format is either x*(N*M)Y

or exp(x+y*N*M), where N is the number of spacecraft and M is the

mass in kilograms. Note DLH and RLH are given in thousands of

hours.

NRLH = DLH - RLH

Structure and Devices

DLH = 1.626"(2"285)A0.9046 = 947.1

RLH = 1.399"(2"285)A0.7445 = 264.0

NRLH = 683.1

Inheritance

Class

3

Thermal Control. Cabling & Pyrotechnics

DLH = exp*(4.2702 + 0.00608"4"30) = 148.4

RLH = 3.731"(4"30)A0.6082 = 68.6

NRLH = 79.8 3

Propulsion

DLH = 56.1878"(4"120)A0.4166 = 735.6

RLH = 1.0"(4"120)_0.9011 =260.7

NRLH = 474.9 3

Attitude & Articulation Control

DLH = 21.328"(4"49)A0.7230 = 968.8

RLH = 1.932"(4"49) = 378.7

NRLH = 590.1 2

31
T¢lecommunications

DLH = 4.471"(4"20)^1.1306 = 633.9

RLH = 1.626"(4"20)A1.1885 = 297.1

NRLH = 336.8

Antennas

DLH = 6.093"(4"5.1)^1.1348 = 186.6

RLH = 3.339"(4"5.1) = 68.1

NRLH = 118.5

Command & Data Handlin_

DLH = exp (4.2605 + 0.02414*4*49.7) = 8600.1

RLH = exp (2.8679 + 0.02726*4*49.7) = 3972.6

NRLH = 4627.5

RTG Power

DLH = 65.300*(4*60)^0.3554 = 458.0

RLH = 7.88"(4"60)^0.7150 = 396.6

NRLH = 61.4

Line-Scan Imagine

DLH = 10.069"(4"36.5)^1.2570 = 5291.5

RLH = 1.989"(4"36.5)^1.4089 = 2228.4

NRLH = 3063.1

Particle & Field Instruments

DLH = 25.948"(4"39.0)^0.7215 = 991.8

RLH = 0.790"(4"39.0)^1.3976 = 917.8

NRLH = 74.0

Remote Sensin,, Instruments

v

DLH = 25.948"(4"44.5)A0.5990 = 578.2

RLH = 0.790*(4*44.5)^0.8393 = 61.2

NRLH = 517.0

2

2

3

3

2

2

2

32
Section 2:

This section analyzes the Development Project - Support

Functions and the Flight Project. PPL is in units of pixels/line. MD is

the mission duration in months and ED is the encounter duration in

months.

PPL = 1024

MD = 151.2

ED = 4.0

EDLH(hardware) = 19540

NRLH = DIM

System Support & Ground Equipment

DLH = 0.36172(ZDLH)^0.9815 = 5887.3

Launch + 30 Days Operations & Ground Software

DLH = .09808(EDLH) = 1916.5

Imaging Data Development

DLH = 0.00124(PPL)^1.629 = 99.4

Science Data Development

DLH = 27.836(non-imaging science mass)A0.3389 = 124.7

Program Management/MA&E

DLH = 0.10097 (EDLH all categories)A0.9670 = 602.5

Flight Operations

DLH = (EDLH/3100)A0.6*(10.7*MD + 27.0*ED) = 5208.8

D_ta Anlalysis

DLH = 0.425*(DLH Flight Operations) = 2213.7

33
Section 3:

Total Costing:

This section incorporates inheritance into the costing. Costing for

class 2 =I.00(RLH) + 0.2(NRLH). Costing for class 3 = 1.00(RLH)

+0.75(NRLH). Since both equations represent labor hours, they must

be converted to dollars.

LH = labor hours = (1.0-Z)*NRLH + RLH

Z = percent cost reduction

LC = labor cost

TC = total cost

Cost Category LH

Structure & Devices

Thermal Control, Cabling &

Pyrotechnics

Propulsion

Attitude & Articulation Control

Telecommunications

Antennas

Command & Data Handling

RTG Power

Line-Scan Imaging

Particle and Field Instruments

Remote Sensing Instruments

System Support & Ground Eq

Launch+30 days Ops & Ground S/W

Image Data Development

Science Data Development

Flight Operations

Data Analysis

Totals

776.3

128.4

616.9

496.7

364.5

91.8

7443.2

442.7

2841.0

932.6

164.6

5887.3

1916.5

99.4

124.7

5208.8

2213.7

29749.1

26975.0

4369.8

23511.7

17671.9

12205.8

3169.1

227894.7

13375.4

108225.8

33624.8

5760.3

191053.5

65969.6

3565.5

6344.0

176571.4

79155.3

999443.6

Total Cost of the Copernicus mission: $ 999,443,600

34
Equations Pertaining to MMPCTC = total cost = (100%-Z) NRC + RC

see costing section of appendix for individual component equations.

Final Trajectory Orbital ElementsOn the following page is an excerpt containing the orbit elements

for the final design trajectory. This contains various data, including

but not limited to flight time, A V required, and the Cartesian

coordinates of significant encounters.

35
L_._'•. -.,. ""....... .........._,

•.,..... =.= =._ ........... _......... ;'.".'? ;12:'/"

C-_ _', /. _ i"--"."_c=- ." i_:',x:-,_ _.'_

_ .,"r.--_ _. "2) _..---'t"_ -_ "-_ _" "_ -'r"i _F"

.... E':7"?

C o;;__;;_';; ":..;;;: ; ; " .........

.... " -''"" "-_ _:;7

I

.L

'7

!, W"_

"_"""_ .....""'..... : _'?"7.,?

,.,'__?;'":- ....• , "!,..,., .. "

,._ . ;...'.•.......

...._"-'" '=;._;;'7"................................

<; ........... '_ ,2,,..-,

................... _ ':. 7 --,;

...... _-"...."• _ "_ .... -'"_ ".'! ;;;7 '-:,7:"7;;_

_ :'= I '- z, '"_• _"_ ' ,'c" _ _. " . -- .........

..... , ,... :......

/ .......

/.i: _'-" ;.......... =;..... C'....... ': ,-,=,.........

,_iI-'_ "w _ .."p.,i C'.',I_.,, p,. T"_ _-, ,-_ r', ,., ,", _', ,- _',. _-,e"_,_/'-'.",,'_'r_

. ...,

=-7 T":+

D

'.:_; 7.'-_

• ,,......,....+,.. . . ..'.... _-:"7 .......................................

;7: :) :? 7

ORIGINAL PAGE IS

OF POOR QUALITY 36
1

1

Appendix 2A. References

"Planet", McGraw-Hill Encyclopedia, V.13, New York:

McGraw-Hill, 1987. pp. 566-569.

2. Kepes, Paul, mulimp files,1990. Files are furnished

at the end of report, enc.

"Jupiter",McGraw-Hill Encyclopedia. V.9, New York:

McGraw-Hill, 1987. pp.482-488.

4. Fink, Donald E.,"Specifications", Aviation Week,

March 19, 1990, pp. 163-64.

o Binzel, Richard P., "Pluto-Charon Mutual Events",

Geophysical Research Letters, Vol.16, No. 11,

November, 1989, pp.1205-1208.

o Dobrovolskis, Anthony R., "Dynamics of Pluto and

Charon", Geophysical Research Letters, Vol. 16,

No. 11, November, 1989, pp.1217-1220.

37
COMMAND,

CONTROL,

AND

COMMUNICATIONS

38
Computer Control

Copernicus, like most spacecraft, must perform a variety of

functions at precise times with unerring accuracy. In order to do

this, an on-board computer system is necessary. The computer

system must control three main areas, the attitude and articulation

subsystem (AACS), the flight data subsystem (FDS), and the computer

control subsystem (CCS). A schematic layout of the computer system

is shown in Appendix 3A. The computer will be made of three

separate, freestanding but interacting computers, controlling the

three areas mentioned. This system is modeled after the system on

board the Voyager spacecraft.

The FDS computer is responsible for all of the flight data

received during the lifetime of the spacecraft. All of the data from

the science platform as well as all the periodic status reports of the

spacecraft are fed into this computer, where it is assimilated,

reduced and passed on. The FDS computer will be a 16 bit x 8192

word computer, as on the Voyager, and will interact with the rest of

the computer system as well as the science platform and most other

instruments for status reports.

The AACS computer is responsible for keeping Copernicus going

in the right direction, with the correct orientation in space. All

tracking data is fed into the AACS computer and it decides if a

readjustment burn is necessary to correct its trajectory. Every

reorientation of the spacecraft, to allow burns or communications, is

timed and the AACS computer knows when to command the burns

and precisely how long to burn. The AACS computer will be an 18

bit x 4096 word computer. This provides ample room for all of its

programming needs.

The CCS computer is also an 18 bit x 4096 word computer. Most

of the permanently stored programs are kept in this computer. If

necessary it can completely reprogram both the AACS and the FDS.

This provides a vital redundancy factor for the spacecraft computer

system. Should the CCS need reprogramming, that would need to be

done from Earth. All information to be sent to Earth and all incoming

information from Earth goes through the CCS computer before

39
moving on to the other computer subsystems, the antenna, or other

areas of the spacecraft.

The three components of the overall computer system interact

fully and all feed into a central storage unit, as shown in Appendix

3A. The data storage unit has a 400 kilobits per second(kbps) record

rate, which will be able to handle all of the incoming data from the

various computer subsystems. It also has five different playback

rates, 100, 50, 25, 12.5, and 6.25 kbps. This wide range will handle

all of the needs of the computers, the science platform and

telecommunications.

Communications

Communications back and forth between Copernicus and Earth is

essential for proper mission accomplishment. Copernicus needs to

relay information such as status reports, scientific data and imagery

back to Earth, while the command center on Earth needs to be able to

send commands to the spacecraft to have it perform certain functions

such as execute a burn, change course or take a picture. While most

of the necessary commands for Copernicus will be stored in the

computer system it is still necessary for communications to be able

to reach the spacecraft.

An antenna is the instrument used to perform the necessary

transmission and collection of data. Copernicus' antenna is a

standard parabolic dish that focuses the radio waves it intercepts to

a central receiving unit, or broadcasts the radio waves onto the dish

which sends them back to Earth.

There are two general radio wave frequencies used in deep

space telecommunications, S-Band "and X-Band. The X-Band is

generally preferred due to its higher frequencies, which have less

interference problems, and it will be used for Copernicus. The XBand uplink (Earth to space) frequency is 7.161 GHz while the

downlink (space to Earth) is 8.414 GHz. There are many factors that

affect the transmission energy before it reaches its destination.

These factors axe summed up by the equation in Appendix 3B. Most

4O
of these factors are losses that reduce the energy from transmission

to reception.

The size of the antenna is the driving factor in the calculation of

necessary power. Large antenna sizes have larger gains, so less

power is needed to achieve a required receiving power. Our antenna

has a significant mass and keeping the mass to a minimum is

important, so we can not allow our antenna to become too large.

Another factor involved in the sizing of the antenna is the fact that it

must fit within our launch vehicle. This means that the antenna

must either be kept small or be collapsible, and much more

complicated. In order to keep the configuration of Copernicus simple

and less costly a solid antenna was chosen. It will be 3.7 meters in

diameter. This provides Copernicus with a small, lightweight

antenna that fits within the launch vehicle but is still capable of

making necessary transmissions with little energy ( app. 25 W).

The positioning of the antenna is vital in mission

accomplishment. The antenna must point towards Earth if

communications between Copernicus and Earth are to occur.

Generally, though, the propulsion for the spacecraft points out of the

back of the spacecraft, towards Earth. The antenna and the

propulsion package will be on opposite ends of the spacecraft. For

most of the beginning of the voyage, the antenna will be useless

because of the fact that Copernicus will be in its burn stage. After

the primary burn stage is complete, Copernicus will rotate 180 o,

allowing full communications. During the flight, if a burn using the

main propulsion is needed, Copernicus must again be rotated 180o.

41
Appendix 3A. Control Flowchart

c¢_

Main Control Communications

AACS

Attitude Control

Tracking

Gyros

S E

FDS

.= Science Data

w

Status Reports

42
B

E

B

E

g

E

E,

E

B

E

E

Appendix 3B. Communications Equation

PR=PTLTC_rrLTI, LsLALpLRpGRLR

PR=Power Received

PT=Power Transmitted

LT=System Losses in Transmitter

GT=Transmitting Antenna Gain

LTp=Pointing Loss of Transmitter

Ls=Free Space Losses

LA=Atmospheric Attenuation

Lp=Polarization Loss Between Antennas

LRp=Pointing Loss of Receiver

GR=Receiving Antenna Gain

LR=System Losses in Receiver

43
Appendix 3C. References

1. Deep Space Telecommunications Systems Engineering

JPL (Joseph H. Yuen) , 1983, Plenum Press, New York & London.

o Draper, R. F. (1988). The Mariner Mark II Program.

Paper AIAA-88-0067, presented at the AIAA 26 th Aerospace

Sciences Meeting, Reno, Nevada, January 11-14, 1988.

J Draper, R.F. (1984). The Mariner Mark II Program.

Paper AIAA-84-0214, presented at the AIAA 22 nd Aerospace

Sciences Meeting, Reno, Nevada, January 9-12, 1984.

4. Heacoek, R.L. (1980). The Voyager Spacecraft.

Paper Proc Instn Mech Engrs Vol 194

44
POWER

AND

PROPULSION

45
Overview

The power system aboard the vehicle utilizes inherently

reliable components. Only materials and techniques available

before 1999 are to be used in the final fabrication of the

system. The system design lifetime is sufficient to carry out the

mission, allowing for a reasonable safety margin. Under normal

mission conditions, the power system is fully autonomous. If

necessary, new commands can be transmitted from the ground

station on Earth. Performance, simplicity, and low weight and

cost are stressed in design tradeoffs.

The main power source is a Modular Isotopic Thermoelectric

Generator (MITG). With the flyby of several planets, the power

requirements will change with respect to the mission timeline.

The modularity of this component makes it ideal for use in this

mission. Releasing power in small scaled amounts, this unit

efficiently meets the power needs of the spacecraft at all times

during the mission.

There exist socio-ecological problems in the use of the MITG,

problems shared with all isotopic thermoelectric generators.

Containing plutonium oxide, debris from these units would be

extremely dangerous in the event of launch mishap. These are

legitimate concerns and have been taken into consideration of

the overall design. For a mission of this duration, however, it is

infeasible to incorporate any other type of system.

46
Table 4-A. Power Requirements

system/component power requirement

AACS 40 W

Science 130 W

Structure

thermal control

pyrotechnics

19.6 W

2.4 W

CCC

computer 24.7 W

data storage 23.2 W

antenna 25 W

Power 25.2 W

Total: 289.9 W

The maximum power required by the system is

approximately 290 W. The total power supplied by the MITG

is approximately 310 W, sufficient for the load requirements.

The maximum power levels will only be reached during

planetary flyby. Here the bulk of the scientific instrumentation

will consume approximately 60 W of power. The imaging

equipment will only be utilized at the encounter with Pluto,

requiring an additional 70 W. The modification of power

supplied will be autonomously controled by the computer.

47
P

0

W

E

R

(w)

290

220

160

I I

JupiWr Saturn

PLANETARY FLYB Y ENCOUNTER S

Figure 4.A.

The earlier planetary encounters require power increases for

only a few days centered about the flyby date. In the case of

Pluto, the imaging process requires weeks of the increased

power level. An insigni_ficant power of 2.4 W is needed for

pyrotechnics at separation of the vehicle from the upper stage.

48
Component Selection

The MITG design was conceived by Fairchild Space and

Electronics Company. They have developed several unit sizes

ranging from output levels of 260 W to approximately 300 W.

Satisfying the power requirement for the spacecraft , the 13

slice generator has been selected.

A redundant circuit design for both the dual busbars and

network has been selected to decrease the chances of failure

due to micrometeorite impact. Parallel fuses are incorporated

on each load to provide redundancy. The electric circuit is

located outside the generator housing, minimizing the

probability of shorts-to-ground problems. Incorporating fieldcancelling circuit modules, scientific instrumentation on the

spacecraft will not be affected by induced magnetic fields from

the MITG.

The generator consists of 13 independent slices each

supplying approximately 24 W at 28 V. Each thermoelectric

slice contains four plutonium oxide pellets supplying a total of

250 W of thermal power. A series of eight thermoelectric

modules per slice convert the thermal power, given off by the

fuel pellets, into electric power for the spacecraft. The

plutonium oxide is contained in an iridium clad surrounded by

an impact shell. Thermal insulation, consisting of carbon

bonded carbon fibers, protects the fuel pellets from under or

over-heating. The whole assembly is protected by an aeroshell,

designed to maintain its structural integrity at extremely high

temperatures. This design uses four radiator fins situated at

the corners of the unit, optimizing heat dissipation as well as

weight.

49
25 W

IL

IF"

M

I

T

G

I

POWER

REGULATOR

I POWER

DISTRIBUTION

w/Main Bus

Protection

22 W

/

THERMAL

CONTROL

40 W

/

130 W

AACS CCC SCIENCE

Figure 4-B. Power Breakdown

5O
_= ", _ J

/ I

Fig. 4-C. Modular

\

I

Isotope Thermoelectric Generator

Pu FUEL

PELLET

GENERATORJ

HOUSING

INSULATION

IRIDIUM CLAD

RADIATOR FIN

IMPACT

SHELL

AEROSHELL

Fig. 4-D. Thermoelectric Slice

51
8

G

0

E

D

Overview

The propulsion system for the vehicle is characterized by

simplicity and reliability. Components incorporated in the system

have been flight tested extensively, meeting with proposal

requirements on availability before the year 1999. The propulsion

unit as well as the fuel storage have design lifetimes sufficient to

carry out the mission, allowing for use of the thrusters for

unexpected mid-course maneuvers. The system relies on

autonomous control by the onboard computer. Performance, weight,

and cost have been optimized in design tradeoffs.

The fuel used in this system is augmented hydrazine. Similar to

conventional hydrazine, it is space storable for long periods of time.

Considering the longevity of this mission, storability is essential.

Because it is a monopropellant fuel, oxidation systems are not

needed, lowering cost and weight. Generally systems of this type are

capable of specific impulses of 200 to 250 seconds. With the use of

augmented hydrazine, values of 300 seconds specific impulse can be

obtained. Advantages of augmented hydrazine include low plume

contamination and no surface contamination, problems which could

interfere with the normal operation of the spacecraft and scientific

instrumentation on board.

The main thrusters will burn twice during the mission. These two

burns will provide the spacecraft with a total AV of 6.1 km / s. The

first burn required is a small mid-course impulse, taking place

approximately ten months after launch. The next burn is at Jupiter

flyby, approximately two years later. This schedule provides for a

smaller probability of error in the propulsion system since all the

major burns occur in the first three years. The remaining amount of

fuel, used by the attitude and articulation thrusters, will be

approximately 5 % that of the initial supply.

52
E

0

VALVE BODY

\

I \

I

INLET FILTER

INJECTOR

DISTRIBUTION

INJECTOR FEED TUBE

EXHAUST

UPPER LOWER

CATALYST CATALYST

BED BED

Figure 4-E. Hydrazine Thruster

The fuel storage tank is characteristic of the bladder design,

eliminating the need for a pressurizing system. As the fuel is

consumed the bladder folds in on itself providing the thrusters with

a steady supply of fuel during burns. Since the attitude and

articulation thrusters also use the hydrazine fuel, the storage tank

can be shared between the two systems. Fuel from the tank travels

through an inlet filter, which removes all foreign particles from the

fuel stream. From there, it is driven through an injector feed tube

and into the injector distribution element. The fuel then passes

through the catalyst bed where it is ignited chemically. Heaters are

situated around the catalyst bed for the chemical reaction to be

carried out properly. Exhaust gasses then escape out of the nozzle

providing the spacecraft with the necessary thrust.

53
Appendix 4A. Equation for Propulsion Subsystem

A V = go Isp in (mi/ mr)

A V = change in velocity

go = constant for gravity

Isp = specific impulse

mi = initial mass

mf = final mass

54
Appendix 4B. References

"Thermoelectric power - the road behind and ahead _

Author: Hall, W. C.

10th International Telecommunications Energy Conference Vol.3

"Rover Power System for Mars Sample Return"

Author: Bents, D. F.

Journal of Space Nuclear Power

"Electrical Power System Architectures

Vehicles"

Authors: Dige, Mark W.; Harold, Neal C ;

Boeing Advanced Systems

for Future Aerospace

Mehdi, Ishaque S.

"Modular Isotopic Thermoelectric Generator _

Author: Schock, A.

"Advanced Propulsion Systems for Space Application"

Author: Knapp, David Edwin

AAE 241 Class Notes - Power Subsystem

AAE 241 Class Notes ~ Propulsion Subsystem

55
SCIENCE

INSTRUMENTATION

56
°';

t  ̧.

d

U

!i i

Introduction

This section describes the scientific subsection of project

Copernicus. This includes a science time line, the planned

experimentation, and the equipment needed to complete the mission.

The selection of experiments was based on present day scientific

objectives for information gathering of the outer planets. Individual

instrument systems were compared and selections were made based

on experimental need. In addition, the requirements and constraints

of NASA's Request For Proposal (RFP) were obeyed.

Voyage to Pluto and Charon

The long voyage to Pluto and Charon will allow an excellent

opportunity for Copernicus to gather information on the galaxy. This

time will not be wasted. During every phase of the journey,

experimentation will take place.

Earth-Jupiter Cruise Phase

After initial Earth orbit and spacecraft deployment have been

established, the science mission will begin in earnest. Once out of

Earth orbit the scientific equipment will be tested and calibrated

through relay with mission scientists on Earth. Later in the journey

such fine tuning will not be possible. Copernicus will spend the

majority of its time in interplanetary space, at these times science

will act in cruise mode. During cruise phases, fields and particles

experiments will be employed. Distant stars will be targeted for

observation and data recording. Information will be gathered and

relayed to Earth approximately every 0.5 AU.

57
E

Jupiter Encounter Phase

As Jupiter nears the instrumentation and experimentation will

convert to encounter mode. The scan platform will be turned to

focus directly on Jupiter. Approximately 80 days and 80,000,000 km

before the closest approach to Jupiter, Copernicus's imaging

equipment will come to life. Over the next seven weeks, the narrow

angle camera will take visual information of the whole planet. A

series of color filters on the camera will also be employed. At this

time, the infrared and ultraviolet spectrometers along with the

photopolarimeter will be taking whole planet data.

As Copernicus approaches 30,000,000 km from closest

approach, the transmitter will begin sending information at

encounter data rate. At this time, the wide angle camera and its

color filters will be engaged. The fields and particles experiments

will also be placed in encounter mode. Specifically, they will

investigate the transition from the region of space dominated by the

solar wind to that of Jupiter's magnetosphere.

As closest approach nears, the equipment on the scan platform

will take advantage of the change in phase angle, from low phase

angles to high, to observe any differences in information due to the

phase angle change. During Jupiter pass by, the Earth will be

eclipsed from Copernicus which will allow an excellent opportunity

for mission scientists observe the effects of the Jovian atmosphere on

the communications signal. This radio science information could be

used to draw conclusions about the composition and height of the

Jovian atmosphere.

As Copernicus leaves it will pass through Jupiters shadow

which will allow ultraviolet inspection of the atmospheric upper

layer composition. Also, long exposure imaging of Jupiters night side

will take place. As the probe continues out the fields and particles

experiments will investigate the extended tail of the magnetosphere.

Transmission will return to cruise data rate 40 days after closest

approach.

58
Jupiter-Saturn Cruise Phase

Upon entering the Jupiter-Saturn cruise phase, science

investigations will return to primarily fields and particles. Special

attention will be paid to the gradual changes in the character and

temperature of the solar wind. Particles experiments will emphasize

the cosmic ray environment. During this phase, the annual solar

conjunctions allow radio science the opportunity to investigate the

solar corona. As communication signals transverse the solar corona

mission scientists can measure the coronal electron density.

Saturn Encounter Phase

The Saturn encounter will progress as did the Jupiter

encounter. The only difference being the emphasis on Saturns rings.

Imaging will begin 80 days out, fields and particles experiments and

transmission rates begin encounter mode 30 days out, and

Copernicus returns to cruise mode 40 days after closest approach.

The information gathered from the Jupiter and Saturn encounters

can be compared to data obtained from the Voyager missions. Any

differences found could be very useful in understanding our

changing planets and galaxy 4.

Saturn-Pluto Cruise Phase

In the final interplanetary cruise phase Copernicus will

investigate the proton component in the distant solar wind plasma.

It will also measure the intensity, composition, and differential

energy spectrum of galactic cosmic rays. These experiments are very

important, as no other spacecraft has taken this final route.

The power and data rate requirements of the science

subsection are shown in time line format in Figure 5A. This clearly

portrays the distinct peaks of power use and transmission

requirements during the planetary encounters. The power

capabilities and communication needs are adequately met by the

Copernicus spacecraft. Figure 5B indicates the individual

instruments used in each phase of the mission. The instruments

were selected for each phase to maximize the data gathering and to

minimize the power drawn and the data transmitted.

59
G

U

G

B

0

0

E_

2.0

8

1.5

6

_:::_ 1.0

,_

o 0.5

2

0

w_.ns_

m

Narrow

Infrared intedeacme_-r

_otopot,_etcr

Ultraviolet

Lowenergycharsed

High_-ersydmled

Magne_me_r

Pl_xm wave de_t_r

Commuai_

mas

0

b

b

, i I , , , I ,

EncounterTimeline

Figure5A Power and Data TransmissionT'mleline

t , I ,

%

• @ • • • • • • •

, , I , , , I , , , I ,

EncounterTimeline

Figure5B InstrumentUse Timeline

6O
Pluto and Charon Encounter

The mission culminates with the investigation of Pluto and its

satellite Charon. Scientific objectives for the two bodies were based

on those from the National Academy of Sciences objectives for the

outer planets 5. Experiments and investigations specific to Pluto and

Charon were developed that would fulfill the needs of the scientific

community. These experiments in approximate order of importance

can be seen in Table 5A 2,6,7. Many of the investigations have specific

subexperiments.

Experiments in Approximate Order of Importance

1. Total Mass and Density

• Map the surface albedo distribution

• Investigate ice to rock ratio

• Investigate composition and hydration state

2. Radius and Oblateness

• Find global maps of Pluto and Charon

• Investigate hydrostatic shape changes

• Map solid body shapes

3. Atmospheric Composition

• Investigate atmospheric induced

limb darkening effects

4. Gravitational Harmonic Coefficients

5. Shape and Strength of magnetic Field at Several Radii

6. Pattern and Magnitude of Heat Flux, Surface Temperature,

and Heat Balance at Various Latitudes and Phase Angles

7. Shape and Intensity of the Tail of the Magnetosphere

or of the Cavity in the Solar Wind

8. Local Anomalies

• Investigate possible dark spots and rings

Table 5A Experimental Listing

61
The experiments will proceed in a similar manner to the

encounters of Jupiter and Saturn. However, because the system

being investigated is a two body system, care will need to be taken

with respect to time management. The scan platform will need to be

rotated to allow adequate time to gather data from both Pluto and

Charon.

As Copernicus approaches Pluto and Charon, the imaging

equipment will begin the investigations of radius and oblateness. It

will begin to compile images that will be used to create global and

solid body maps, and to investigate any hydrostatic shape changes.

These maps will be used to help determine the radius and oblateness

of both bodies. As the probe nears, the infrared interferometer

spectrometer will be used to investigate thermal emissions,

composition of thermal structure, and heat balances. This data will

be collected over a variety of phase angles. The information, along

with the imaging data will help to map the surface albedo

distribution, investigate the ice to rock ratio, find the pattern and

magnitude of heat flux, surface temperature, and heat balance at

various latitudes and phase angles of both Pluto and Charon. While

still on approach, Copernicus will accumulate data with its

magnetometer. Information from the magnetometer will aid in

determining each bodies gravitational harmonic coefficients and the

shape and strength of their magnetic fields at several radii. The

photopolarimeter will investigate the physical and chemical

properties of Pluto and Charon. This information, along with data

from the infrared interferometer spectrometer and the imaging

equipment, will help to determine the composition, mass, and density

of both bodies.

As the spacecraft passes through its closest approach, the

particles experiments will convert to encounter mode. In this mode

they can gather a variety of important information. The high energy

particle detector will measure electrons and cosmic rays, while the

low energy particle detector investigates particles in the planetary

magnetosphere. The plasma particle detector will determine plasma

flow direction and the plasma wave detector will study the wave and

particle interaction in the dynamics of the magnetosphere. All this

62
E

E

E

E

ntr

_-_ b

E

U

E

information will be used to model the shape and intensity of the tail

of the magnetosphere or of the cavity in the solar wind.

While the probe is eclipsed from Earth, additional

investigations will be made. The imaging equipment will focus on

Plutos limb and terminator region. Data acquired can be used to

determine the atmospheric induced limb darkening effects.

Communications tracking of the probe can aid in finding Plutos

gravitational harmonic coefficients and the strength of its magnetic

field.

As Copernicus sails into the outer galaxy its investigations will

not end. Possibly it could investigate the heliopause. It wilt continue

to send data from our galaxy back to Earth.

Equipment Selection

The design features a wide variety of imaging, spectroscopy,

and fields and particles instruments. All equipment was selected

from existing hardware used on the Cassini, Galileo, and Voyager

missions. This was done to minimize cost while keeping a high level

of information accuracy and reliability.

Imaging Science Subsystem (ISS)

The Copernicus probe will encounter a wide variety of targets

and range of observing distances. Therefore, two separate cameras

will be used in the ISS, a Narrow Angle Camera (NAC) and a Wide

Angle Camera (WAC). In this way, Copernicus can provide two

different scales of image resolution and coverage.

The two cameras are framing Charge Coupled Device (CCD)

imagers. The charge couple device design is a square array of

1024 x 1024 pixels, each pixel is 12 ttmeters on a side. They differ

primarily in the design of the optics: the NAC has a focal length of

2000 mm and the WAC has a focal length of 250 mm. Both cameras

have a focal plane shutter of the Voyager/Galileo type, and a twowheel filter changing mechanism derived from the Hubble Space

Telescope. Both cameras have deployable dust covers. To minimize

mass, power, and cost, the two cameras will not be completely

63
independent - they will share a common electronics module. This

module services both cameras, and contains the digital part of the

video signal chain, power supplies, mechanism drivers, command and

control logic, and the digital data compressor3.

Key parameters of the ISS:

Narrow Angle

Camera Type

Optics Type

Focal Length

Focal Ratio

Resolution per pixel

Field of View

Spectral Range

Spectral Filters

Heater Unit

Framing CCD

Ritchey-Chretien

2000 mm

f/10.5

6 I_rad

0.35 ° square

200-1100 nm

22

Strip heaters

Framing CCD

Refractor

250 mm

f/4.0

48 I_rad

2.8 ° square

350-1100 nm

14

Strip heaters

Infrared Interferometer Spectrometer (IIS)

This instrument consists of an infrared radiation telescope, two

Michelson interferometers for evaluating spectral data, and a

radiometer for measuring total body reflection. The IIS will be used

to measure planetary thermal emissions, surface composition, and

thermal structure. It will accomplish this by measuring reflected

solar radiation and heat balancesl, 4.

Photopolarimeter

The photopolarimeter gathers information on surfaces or

particles by observing how they scatter light. To accomplish this the

photopolarimeter must take measurements over a variety of phase

angles. This data can be evaluated to find the physical and chemical

properties of planetary atmospheres and surfaces. The intensity and

polarization of light are measured in 10 narrow bands from 0.41-

0.945 microns, including areas where methane and ammonia

strongly absorb radiationl, 4.

Ultraviolet Spectrometer (UVS)

64
The ultraviolet spectrometer operates in two distinct modes:

airglow and solar occultation. During Copernicus's cruise phases, the

UVS will operate in airglow mode. It will observe the sources of

extreme ultraviolet radiation in the galaxy. As the probe enters an

encounter phase and passes by a planet, the ultraviolet spectrometer

will convert to solar occultation mode. In this mode the instrument

will study solar light and the effects a passing planets atmosphere

has on it. The UVS covers a 0.115-0.43 micron spectrum and views

with a 0.1 ° slit width. The ultraviolet spectrometer can detect

nitrogen, sulfur, and atomic hydrogen and oxygen. Microprocessor

control provides flexibility. The UVS can fix at one wavelength and

look for intensity changes during a scan, or it can rapidly step

through wavelengths for a full spectrum over a broader area - or

some combination in between 1,4.

Particles Investigations

The particles studies consist of three distinct instrument

investigations. They are a Low Energy Charged Particle (LECP)

detector, a High Energy Charged Particle (HECP) detector, and a

Plasma Particle (PP) detector. The LECP detector operates with two

objectives: measure particles in planetary magnetosphere and to

detect low energy charged particles in interstellar space. It

accomplishes its objectives by measuring particle source,

composition, energy spectra, flux intensity, and favored particle

direction. The HECP detector is similar to the low energy charged

particle detector, however it measures particles by charge, mass,

energy, and arrival direction. The LECP and HECP work with a

combined range of 0.020-55 million electron volts for ions and

0.015-11 million electron volts for electrons.

The plasma particle detector consists of two Faraday cup

plasma sensors and three mass spectrometers. Its objective is

measuring the plasma in the solar wind and in planetary

magnetospheres. It is also responsible for finding the plasma flow

direction. The PP detector studies plasma by detecting its velocity,

density, and pressure. This device measures the energy range of

electrons and positive ions from 1.2-50,400 electron volts. The

65
Faraday cup plasma sensors collect the plasma data, while the three

mass spectrometers are included to identify the composition of

ion s 1,4.

Fields Investigations

The instruments that fall under the fields category are the

magnetometers and the plasma wave detector. The magnetic fields

investigations employs four magnetometers. This investigation uses

two sets of two triaxial fluxgate magnetometers. One set is of low

field, the other high field. These magnetometers measure planetary

magnetic fields. They measure with a range of 0.00032-0.16384

gauss.

The plasma wave detector will be used to study wave and

particle interaction in the dynamics of a planets magnetosphere. The

detector measures changes in electric and magnetic fields. The

electric and magnetic fields can be measured separately over ranges

of 5 Hz. to 5.6 MHz. and 5 Hz. to 160 KHz., respectively 1,4.

Table 5B shows the scientific mission at Pluto/Charon of each

instrument Copernicus will be carrying. All equipment will be

heated with a combination of strip heaters and passive

athermalization with invar and aluminum structures.

Instrument Layout

Instruments will reside in one of three locations aboard the

spacecraft. The magnetometer boom, the scan platform, or the scan

platform boom. The scan platform and its boom, along with the

magnetometer boom were located so as to maximize their distance

from each other and from the Radio Isotope Thermal Electric

Generator (RTG).

Scan Platform

The scan platform will house the instruments that specifically

need to be pointing at the target they are investigating. It will be

extended out from the Copernicus by a folding boom. The platform

itself will have two axis of freedom about which to rotate. This will

66
B

B

E

B

E

E

B

;]

E

EQUIPMENT

Imaging

INVESTIGATION CONCERNS

Radius

Oblateness

Global maps

Solid body maps

Limb darkening

Map surface albedo

Ice to rock ratio

Mass

Density

Terminator region

Interferometer

Spectrometer

Thermal emissions Map surface albedo

Composition of thermal structure Ice to rock ratio

Heat balances Surface temperature

Heat flux Composition

Mass Density

Magnetometer Harmonic coefficients Magnetic fields

Photopolarimeter Physical, chemical properties Composition

Mass Density

HECP Detector Measure electrons Measure cosmic rays

Tail of magnetosphere Cavity in the solar wind

LECP Detector Particles in magnetosphere Tail of magnetosphere

Cavity in the solar wind

Plasma Particle Detector Plasma flow direction Tail of magnetosphere

Cavity in the solar wind

Plasma Wave Detector Particle interaction Tail of magnetosphere

Cavity in the solar wind

Ultraviolet Atmospheric composition

Spectrometer

Table 5B Insmanent Investigations

minimize the maneuvering required from the spacecraft. The

instruments on the scan platform include the narrow angle and wide

angle cameras and their electronics, the infrared interferometer

spectrometer, the photopolarimeter, the ultraviolet spectrometer,

and the plasma wave detector. The equipment will be placed

together and bore sighted with the narrow angle camera. By placing

the instruments in a cluster, the strip heaters can serve more than

one instrument, thereby minimizing power use and cost. Because the

67
6_

D

B

C

equipment will be bore sighted on the narrow angle camera, mission

scientists will have an image corresponding to data collected from

the other scan equipment.

Requirements are placed on the movement of the scan platform

by the science instrumentation, specifically the imaging equipment.

The platform can rotate with a maximum slew rate of 0.33 ° per

second. At this rate the instruments with the exception of the

imaging equipment can be accurately used after a settling time of 45

seconds. However, if the cameras are to be employed a settling time

of 288 seconds is required. The equipment on the scan platform also

places a limit to the maximum maneuver rate of the spacecraft. The

maximum allowable maneuver rate of Copernicus while performing

experiments, except imaging is 0.033 ° per second. The maneuver

rate while imaging drops to 0.00972 ° per second. Another

requirement for the scan platform is its pointing accuracy. The

platform must be high precision with pointing accuracy of at least 2

mrad with 1 mrad knowledge and stability of 10 mrad in 0.5 seconds

and 100 mrad in 100 seconds. Figure 5C represents a view of the

scan platform and its equipment3,8.

Wide A%le Camera _ I u [

L t

Camera

Photopolanmeter _ _

Figure 5C •

Imaging Electronics

Ultraviolet

Spectrometer

Infrared

Interferometer

Spectrometer

68
Scan Platform Boom

The scan platform boom is a convenient location to place the

particles instruments. It is away from the spacecraft and allows

"undisturbed flow through of the interstellar environment. The boom

will house the low and high energy charged particle detectors and

the plasma particle detector.

Magnetometer Boom

This 13 meter long boom will remove its low field

magnetometers from interference with the other science equipment.

The magnetometers will be the only instruments placed on this

boom. The high field magnetometers will be place on the boom near

its attachment to the spacecraft. One low field magnetometer will be

located half way down the boom, the other placed at the farthest

end.

Table 5C is a listing each instrument and its mass, power

requirement, data transmission rate, and location on the probe 4,8.

INSTRUMENT

tmaging

Infrared

Interferometer

Spectrometer

Photopolarmeter

Ultraviolet

Spectrometer

LECP Detector

HECP Detector

Plasma Particle

Detector

Magnetometer

MASS

(kg)

POWER DATA LOCATION

0,V) RATE (bps)

36.5 29.0 3850 Scan Platform

18.5 12.0 500 Scan Platform

13.0 13.0 450 Scan Platform

13.0 13.0 450 Scan Platform

9.0 16.0 450 Scan Boom

13.8 16.5 450 Scan Boom

9.9 8.1 450 Scan Boom

4.9 5.8 400 Magnetometer

Boom

Plasma Wave 1.4

Detector

1.6 200 Scan Platform

Table 5C Immanent Data

69
Conclusion

The mission to Pluto and Charon can only be completed cost

effectively by a spacecraft whose science section maximizes accurate

data gathering and the number of target investigations, while

minimizing mass, power consumption, and complexity. The

Copernicus probe meets these requirements.

70
APPENDIX 5A

References

1. Aviation Week & Space Technology, Galileo Represents Peak in

Design Complexity, P77. (October 9, 1989).

2. Croswell, Pluto: Enigma on the Edge of the Solar System, P8-22.

Astronomy (July 1986).

3. Jet Propulsion Laboratory, Vol. VI: Imaging Science Subsystem

Description; Cassini Mission: Saturn Orbiter Proposal Information

Package, P1-8. JPL D-6464 (October 10, 1989).

4. Jet Propulsion Laboratory, Voyager Neptune/Interstellar Mission;

Flight Science Office; Science and Mission Systems Handbook, P1-248.

California Institute of Technology PD 618-178 (1987).

5. National Academy of Sciences, Conference on Outer Planetary

Exploration, Pl-12. (1978).

6. Reynolds, The Interiors of Pluto and Charon: Structure,

Composition, and Implications, P1209-1212. Geophysical Research

Letters Vol. 16 No. 11 (November 1989).

7. Trafton, Pluto's Atmosphere Near Perihelion, P1213-1216.

Geophysical Research Letters Vol. 16 No. 11 (November 1989).

8. Science Working Group, Report of the Comet Rendezvous. NASA

Technical Memorandum 87564 (1985).

71
E

B

E

E

E

E

B

E

E

E]

ATTITUDE

AND

ARTICULATION

CONTROL

72
E

0

8

0

E

0

0

Introduction

The task of the Attitude and Articulation Control System (AACS)

is to control the attitude of the spacecraft. This requires pointing the

high gain antenna toward the Earth and/or Sun, pointing the

trajectory correction thrusters in any direction, providing control

authority during the rocket engine burns, performing science

maneuvers, and pointing the scan platform.

These control requirements are very challenging because of the

complex and time changing parameters the Copernicus will

encounter. Initially, there is the change in mass at separation from

the launch vehicle, and then the changes in mass during mid-course

correction and orbit burns. Propellant slosh is and wobble

amplifications are also factors.

These requirements and the time-varying parameters dictate a

complex set of AACS sensors and actuators controlled by a high

performance computer, and that a great deal of on-board autonomy

be present in the AACS. Also there are weight and power constraints

that put stringent requirements on the electronic components. A

mission objective is to prevent single-point failures from

jeopardizing the mission. This forces redundancy of the critical

components and requires internal fault protection logic to control

that redundancy.

Without doubt, accurate attitude control of the Copernicus is

imperative to mission success. This section describes the attitude

control of the Copernicus spacecraft during the entire mission, giving

detailed descriptions of the components and methods used in

designing the AACS.

Attitude Control Modes

The attitude of the Copernicus is achieved through the use of a

set of celestial sensors, a set of inertial sensors, an onboard digital

computer, and a set of hydrazine thrusters. The Copernicus will be

73
E

E

E]

B

B

E

E

E]

B

D

G

E

E

three-axis stabilized due to the science requirement for a scan

platform and the lower cost compared to a dual spin design. Threeaxis stabilization also permits extended viewing of selected targets,

thus permitting a larger number of individual measurements or a

longer integration time for increased sensitivity per measurement

than can be achieved with a spin stabilized spacecraft unless it has a

de-spun platform.

On account of the length of the mission, the Copernicus must be

able to function autonomously for a large amount of its travel time.

A basic guideline is that the spacecraft (S/C) be able to operate for at

least one week without ground intervention without loss of more

than one science instrument or loss of more than one-half the

engineering telemetry and the S/C must be left in a commandable

state. Therefore it is imperative that the control computer have

various fault detection and correction actions when the S/C

subsystems experience certain failures, and be able to maintain

correct attitude control during these times 6.

A software estimation process has been derived to determine the

best spacecraft position, rate, and acceleration estimates in the

presence of noise and disturbance processes. Based on these

estimates the attitude of the spacecraft is corrected by activating the

appropriate hydrazine thrusters. The algorithm for determining the

best spacecraft position and rate is described in Appendix 6A 1

During cruise, the normal response to a fault is to "safe" the S/C

in a specifically oriented attitude. However, during critical mission

phases, the on-board systems must reconfigure the Copernicus in

such a way as to maximize the probability of completing critical

sequences (such as burn and science maneuvers). To accomplish

various maneuvers necessary in reorienting the Copernicus, a

commanded turn capability is implemented. A turn in any of the

three axes is accomplished by the insertion of a bias in the control

loop during inertial cruise.

74
E

E

B

8

E

E

E

Scanning Platform and Pointing Control

The mounting of a science scan platform at the end of a science

boom permits the physical tie-down of its mass during launch,

provides for mass balancing of the RTG's for spacecraft center of

mass control, and maximizes the unobstructed solid angle through

which the remote sensing instruments can be pointed. This platform

holds all of the science instrumentation and sensor and control

components, which have accurate pointing requirements, thereby

eliminating many sources of error that have existed on prior

spacecraft. Clearly, the pointing performance of this platform is

critical to the success of the mission.

Typical pointing requirements for a high precision scan platform

(HPSP) are shown in Table 6-A. These requirements are primarily

driven from the requirements of the cameras, and apply to each of

Table 6-A. Pointing Requirements

0

E]

High Precision Scan Platform

Inertial Pointing Control

Inertial Pointing Knowledge

Inertial Pointing Stability

(during 0 to 17.5 mrad/sec slew)

Requirements

2.0 mrad (0.11 °)

1.0 mrad (0.06 °)

10 wad/0.5 sec

100 wad/100sec

E

B

E

the two required axes of articulation. These requirements fall well

within the requirements for the entire Copernicus mission. The

dynamics of the platform boom can be excited by both basebody

motion and platform slews. The choice of an appropriate scan

actuator which controls this platform, and compensates for

disturbances, will be described next.

75
D

D

E

E

E

Scan Actuator

A key element in the mission is the high precision scan platform.

On this platform a number of instruments are mounted, including

several cameras and the star tracker and gyro used for S/C attitude

control. Clearly, the pointing performance of this platform is critical

to the success of the mission. The central consideration of a scan

actuator can have an impact on the design of the entire spacecraft.

A direct drive actuator with a platform mounted momentum

compensation wheel is selected for the Copernicus. This actuator is

selected on the basis of net effect on spacecraft mass, required

power, cost, expected pointing performance, necessary control

complexity, suitability to mission, operational considerations, and

ability to accomodate changes in the mission or spacecraft. It is

assumed that all actuators considered met the spacecraft reliability

and lifetime requirements.

Table 6-B compares four models of possible actuators, including a

momentum compensation harmonic drive (MCHD), direct drive,

harmonic drive (I-IDA), and two-motor actuators. It can be seen that

Table 6-B. Scan Actuator Comparison

E

E

0

E

E

Criteria

Reliability

Mass

Total Power Peak/

Steady State

Performance

Heritage

Direct

_riv_

Least Risk

27 KG

8W/6W

1 wad

Galileo

Two-Motor

Acceptable

51 KG

1 7W/1 2W

16 p.rad

Pathfinder

MCHD HDA

Acceptable Unacceptable

50 KG 31 KG

1 1W/8W 1 0W/6W

7 wad

Breadboard

N/A

Halley

Intercept

overall the direct drive actuator is the best choice, with the bonus

that it's been space tested on the Galileo.

The reason for the momentum compensation wheel is that a

savings in attitude control propellant can result in an overall savings

76
E

E

E

E

,E

E

E

E

E

E

E

E

of spacecraft mass for missions requiring a large number of platform

slews, such as Copernicus. Thus when the scan platform accelerates

in azimuth, the motor-mounted wheel with the required inertia ratio

will accelerate in the opposite direction. The elevation axis works

the same way. So ideally the spacecraft body will not sense the

platform articulation disturbance torques.

The direct drive actuator is the simplest of the configurations

considered. It consists of a brushless DC motor mounted at the

gimbal joint. Torque is applied directly by the motor to the platform

and a reaction torque is applied directly to the basebodyS.

Star Tracker

The development of charge-coupled device (CCD) optical sensors

has made it possible to construct high-performance star and target

trackers for spacecraft. They offer high resolution, dimensional

stability, and both geometric and photometric linearity. The

ASTROS-II (Advanced Star/Target Reference Optical Sensor) tracker

currently being developed at the Jet Propulsion Laboratory is

scheduled to be launched on the Comet Rendezvous Asteroid Flyby

mission. This tracker uses the RCA 501 DX CCD, has integral

microprocessors to control the data acquisition, make image position

calculations, and provide an effective interface to the pointing control

computer.

Table 6-C compares available star sensors. The ASTROS-II is

based on the ASTROS built for flight on a series of shuttle-based

ultraviolet astronomy missions. The revised design will be tailored

to requirements of the Copernicus mission. The ASTROS-II has the

following capabilities:

a) Tracks several stars simultaneously for attitude reference

(up to 5 stars per field).

77
E

Table 6-C. Star Tracker

Internal Redundancy

Comparison

Characteristic CS-203 Canopus ASTROS

Mission VRM Voyager Shuttle Copernicus

Field of View 4.6 ° wide 9 ° x 36 ° 2.2 ° x 3.5 ° 1.5°x11.5 °

Drift Rate (°/sec) 0.2-1.0 N/A <.1 <.5

Yes No No Yes

17x24x18

5.5

7

Dimensions (cm) 29xl 3xl I

4.3

4.5

Mass (kg)

Power (w)

50x25x20

2.8

38

* VRM - Venus Radar Mapper

ASTROS-II

25x16x16

8

11

b) Follows rapidly moving , time-varying , extended

targets during a close flyby or rendezvous.

c) Determines the limb position and orientation of a

nearby target.

d) Develops image data for ground-based target searches

during target approach.

e) Tracks both stars and extended targets and provides

optical navigation data for the mission.

f) Mass, power, volume, and environmental compatibility

with the Copernicus mission.

These qualities make the ASTROS-II an optimal choice for the

Copernicus mission. The unit will be internally redundant and

therefore the specifications listed in Table 6-C make it a substantially

better choice than all others. The tracker will be located on the I-IPSP

along with the scientific instrumentation2.

78
Laser Gyro

The attitude of the Copernicus in three-space is measured by a

new technology gyro based on fiber optics, Fiber Optic Rotation

Sensor (FORS). Nearly 100 years ago, it was discovered that light,

along with conventional gyroscopes, could provide gyroscopic

information. The time it takes light to traverse a circular pathway

depends on whether the pathway is stationary or rotating. The time

difference can measure the amount of rotation 7.

The FORS design uses a single 5 mW GaA1As laser to input light,

divided and injected, into both ends of a 3 to 20 km long fiber

waveguide wrapped around an 18 cm coil. After the light has passed

through the fiber waveguide, it is recombined and detected. This

concept is based on the Sagnac interferometer principle. The phase

angle between the two light beams is dependent upon coil rotation

rate, direction, number of turns of the fiber, and area enclosed 3.

There will be two sets of three of these gyros for redundancy.

The use of this type of gyro results in a planetary gyro with ten

times improved drift rate over today's conventional gyros. With the

absence of moving parts, no gas discharge tube, and no short term

wearout mechanisms, the operating lifetime is well within the

mission requirements for Copernicus.

The fabrication processes are relatively inexpensive. The absence

of moving parts and close similarity to electronic microcircuit

fabrication allow this. The recurring cost of these new planetary

gyros is less than one-third of today's conventional gyro cost. The

mass, power, and volume will also be

Table 6-D. Gyro Comparison

Unit Drift Rate(°/sec'_AngulaResolution Power (w) Mass (kg)

2xl 0" 4 0.005 arcsec 1 0 1 0

DRIRU-II 3xl 0" 3 0.05arcsec 22 1 1

CG-1300 Laser 7xl 0" 3 1.4 arcsec 1 8 1 8

Volume (cm 3

16400

16236

5740

79
R

E

B

less than present gyros. Table 6-D compares the FORS and two other

currently available gyros. The entire gyro component will be placed

on the science scan platform for optimal accuracy 3.

Reaction Control System (Thrusters)

The Reaction Control System (RCS) of the Copernicus consists of

twelve 1N thrusters located in four clusters about the center of mass

of the spacecraft, illustrated in Figure 1-C (Structures Section). The

RCS is a monopropellant hydrazine system which has fuel supply

lines running from the main propellant bladder. The thrusters are

similar to the Voyager design and act as couples. They provide

attitude control torques and thrust for small engine maneuvers and

trajectory correction maneuvers (TCM), but the main propulsion

engine provides most of the control thrust during impulse burns and

large maneuvers. The use of four clusters with three thrusters each

provides redundancy, designating main and backup sets of thrusters

which can be used for control about specific axes. An example of the

designated control setup is shown in Figure 6-A and Table 6-E.

The thruster is designed to provide 0.95N thrust and 300s

specific impulse at propellant inlet pressure of 24.6 kgf/cm2a to

meet the requirements for Copernicus. The thruster has a 60:1

expansion ratio conical nozzle. Thrust level is adjusted by controlling

the flow rate of propellant with valves located on the fuel lines. The

amount of propellant reserved for attitude control is estimated to be

about 5% of the total fuel for the mission. This estimate takes into

account the longer duration and therefore many more TCM's which

will take place compared to previous missions, but also realizes the

greater mass of fuel which is being carried for this mission

(compared to other missions) 8.

Thermal design of the thruster cluster uses three catalyst bed

heaters and valve heaters to maintain the catalyst bed above 200 ° C

prior to firing. The cluster is designed to be thermally isolated from

the spacecraft and minimize heat transfer to the cluster or propellant

valve to keep the catalyst bed hot. The thrusters are designed to be

8O
Figure 6-A. Thruster Cluster Configuration

Copernicus Center of Mass

1.914m 10__6 l r 1_ _

+Pitc_h ..._3_ / _

Table 6-E. Thruster Location and Function Matrix

Control Mode 1 2 3 4

+Yaw

-Yaw

+Roll M B M B

-Roll B M B M

+Pitch

-Pitch

Thruster Number

5 6 7 8

B B M M

M M B B

9 10 11 12

M B B M

B M M B

81
E

E

D

E

E

B

E

B

D

E

E

capable of achieving mission requirements even in the case of one

heater failure.

Algorithms within the main computer control the thrusters to

provide three-axis control and to perform closed-loop turns of the

spin axis. Such turns may be required up to four times daily to keep

the high gain antenna pointed toward Earth, and to orient the

spacecraft for TCM's 1.

Conclusion

The Copernicus spacecraft is 3-axis stabilized, using a digital

onboard computer, a set of fiber optic gyros, a star tracker, and

hydrazine thrusters. Attitude control of the spacecraft is based on

measuring spacecraft orientation, estimating spacecraft states, and

actuating the thrusters for attitude correction.

The orientation of Copernicus is measured by FORS. The position

is calculated using the ASTROS-II. A direct drive actuator with

momentum compensation wheel will be used to operate the scan

platform. The attitude of the spacecraft will be adjusted with 1N

thrusters, located on a structure which surrounds the propellant

bladder.

This configuration for the AACS will provide the best control for

the long journey the Copernicus will undertake to Pluto.

82
E

E

E

E

E

E

E

Appendix 6A. Inertial Control Single Step State

State Predictor in Cruise

_tp (K/K) = _tp (K/K-l) + Kp[ Mp(K) - H_p(K/K-1)]

_tp (K+I/K) = tt(K+l,K)2_p (K/K) + P_p(K+I,K)Tp(K)

(15 T2, (5 T2

CK+I,K) = 0 1 AT

0 0 1 / _P(K+I'K)=-_'IP= AT/Jp

The decision to turn the appropriate thruster on at K+I is based on:

Ep(Z+l) = (1 Krp 0) _p(K+I/K)

_p(K/K) is best estimate of spacecraft pitch statres at K given

measurements Mp(K).

._.p(K+I/K) is the best one-step prediction of S/C pitch state based on

Mp(K).

Kp is the Kalman gains.

Tp is the estimate of torque developed by pitch thrusters.

Process is sequentially repeated in real time.

For yaw and roll axes, the subscripts p are changed to y or r.

83
0

0

E

0

8

8

B

8

0

8

0

G

B

8

8

B

0

U

Appendix 6B. References

1. Bahrami, K., "Inertial Attitude Control of Voyager Spacecraft

Using Dry Tuned Rotor Gyro", AIAA Guidance and Control Conference,

Hollywood, CA, August 8-10, 1977 pp. 520-529.

2. Dennison, E.W., Stanton, R.H., Shimada, K., "Astros: High

Performance CCD Tracker for Spacecraft", International Symposium

on Space Technology and Science, 15th, Tokyo, Japan, May 19-23,

1986 Vol.2 pp. 1131-1142.

3. Draper, R.F., "The Mariner Mark II Program", AIAA Aerospace

Sciences Meeting, 22nd, Reno, NV, January 9-12, 1984, 34 pgs.

4. Heacock, Raymond L., "The Voyager Spacecraft", Proceeding of the

Institution of Mechanical Engineering", Vol. 194, 1980, pp. 211-221.

5. Schier, J. Alan, Bell, Charles E., Agronin, Michael, Socha, Michael,

"The Mariner Mark II High Precision Scan Actuator: Results of an

Options Trade Study Covering Point Performance and System-Wide

Impacts", AIAA Navigation, Guidance, and Control Conference,

Minneapolis, MN, August 15-17, 1988, pp. 1053-1060.

6. Stephenson, R. Rhoads, "The Galileo Attitude and Articulation

Control System: A Radiation-Hard, High Precision, State-of-the-Art

Control System", IFAC Automatic Control in Space, Toulouse, France,

1985, pp. 83-89.

7. Stuart, Dr. James P., White, Thomas W., Coffey, Randall E., "Space

System Design and Tradeoffs of Ring-Laser Gyroscope Guidance and

Control", AIAA Aerospace Sciences Meeting, 25th, Jan. 12-15, 1987

8. Yoshiyama, Iwao, Iida, Chisato, Kon-No, Akira, Takiyama, Tetsuma,

Mori, Takashi, Yabuhara, Eiji, "Design and Flight Performance of the

Hydrazine Reaction Control Subsystem for the ETS-III", International

84
Symposium on Space Technology and Science, 13th, Tokyo, Japan,

1984, pp. 221-231.

85
EOE888OEOBO0O0

CONCLUSION

86
E

E

E

E

E

E

E

B

E

Conclusion

This proposal for an unmanned mission to Pluto cans for the

spacecraft Copernicus to be launched on May 4, 2009 on a 12.6 year

journey through the outer Solar System with flybys of Jupiter and

Saturn before it reaches its (possible) final destination of Plutoian

space.

The proposed design adheres to the previously stated mission

requirements and special emphasis was put on optimizing

performance, reliability, and mission cost.

This proposal is only a Phase A design report, but it does provide

the initial research necessary for later more detailed mission

concepts and designs.

87
E0DE

E_EEE

z17_

_3

®o'_ _

o

rl- (.,_ ._i

E
EEEEEEEEEEBE

A

,4

I=0

(,1

lit Q

Iom

II m

Io0a

>,

i._

f6Ee_

I1,,

a.

lira0

I=0m

'ellamQ

I,.,

i._00

e_o

0= "6

(._ |

_EE

._ -, F. -, C :::I"

o 8_i:-_ 0

aE==

"o_ -_'--_.

_o

.g

!

0

",E_S

.__,_ .__

__o

._=

=_

om lira

--w._

_ -.r .5 ,'_

._r ,... 8

_1,_®®

,_-_ "ZO_ "_

i

,.®..=o ®="=

(1.

<

"!!

.0_

E_ w t1_ r- _ E--_

o__

-_.__

_'_.

I'-