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Intrepid: A Mission to Pluto 

Michael Behling 

Donald Buchman 

Andres Marcus 

Stephanie Procopis 

Carl Wassgren 

Sarah Ziemer 

Aerospace Design Class 

AAE 241 

April 24, 1990 

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Table of Contents

List of Figures ................................................ ii

List of Tables .............................. "................... ii

Introduction ................................................... 1

Scientific Instrumetation ................................. 2

References ................................................... 14

Mission Management, Planning, & Costing ....... 15

Appendix A2.1 ............................................ 24

Appendix A2.2 ............................................ 27

Appendix A2.3 ............................................ 28

Appendix A2.4 ............................................ 29

References .................................................. 30

Power and Propulsion Subsystem ...................... 31

Appendix A3 ............................................... 45

References .................................................. 47

Structural Subsystem ....................................... 48

References .................................................. 62

Command, Control, and Communications ........63

References ................................................. 72

Attitude and Articulation Control .....................73

Appendix A6.1 ............................................ 87

Appendix A6.2 ............................................ 87

References ................................................. 88

Conclusion .................................................... 89

Appendix B .................................................... 90

Appendix C .................................................... 92

Appendix D .................................................... g3
List of Figures

1.1

2.1

3.1

3.2

3.3

3.4

3.5

3.6

4.1

4.2

4.3

4.4

4.5

4.6

6.1

6.2

6.3

Layout of Scientific Instrumentation

Intrepid Launch Sequence

Propulsion System Configuration

Inlet Pressure versus Specific Impulse

Schematic of Power Subsystem

Modular RTG Design

Titan IIID/Centaur Launch Energy vs Launch Mass

Titan IVD/Centaur Launch Energy vs Launch Mass

Titan IIID/Centaur

Subsystem Necessities

Structural Design Considerations

Intrepid Spacecraft Flight Configuration - X-Z Plane

Intrepid Spacecraft Flight Configuration - X-Y Plane

Intrepid Spacecraft Flight Configuration - Y-Z Plane

Intrepid Spacecraft in Launch Configuration

HPSP with Attitude and Science Instruments

Bus Thruster Layout

Mission Modes

List of Tables

2.1

2.2

2.3

3.1

3.2

4.1

4.2

4.3

5.1

6.1

Parameters for Trajectory Consideration

Launch Parameters for the Intrepid Spacecraft

Summary of Trajectory Encounters

Breakdown of Power Requirements

Launch Vehicle Information

Method of Achieved Thermal Control

Environmental Considerations

Materials Considered

Performance Parameters for Intrepid Radio System

Comparison of Rate Sensors
Introduction

A proposal for an exploratory spacecraft mission to the Pluto/Charon

system has been written in response to the "Request for Proposal for an

Unmanned Probe to Pluto," or RFP. The RFP lists many design

requirements that must be satisfied by the proposed spacecraft. They are as

follows:

1.

2.

3.

design an unmanned scientific study to Pluto and Charon

mission scrence objectives must be described and justified

optimize performance, weight, and cost of spacecraft in design

tradeoffs

4. launch time: between the years 2000 and 2010

5. design should stress reliability, simplicity, and low cost

6. spacecraft must be able to adapt to whatever environment it may

encounter

7. lifetime of spacecraft must include mission time and a safety

margin

8. nothing in spacecraft's design should preclude it from performing

several missions

9. use existing hardware when possible in design

10. use materials and techniques available by 1999

11. use latest advances in artificial intelligence

12. amount of on-orbit assembly should be identified and minimized

13. identify use of space shuttle if applicable

14. if space shuttle is used, it must apply to NASA standards

15. for cost estimates: assume four spacecraft built, three flight

ready and one for integrated ground test system

Under the guidence of the RFP requirements, the spacecraft Intrepid

was designed. The RFP requirement which was of primary importance is

that to keep costs at a minimum. The less expensive the design is, the more

attractive it would be to those ultimately in control of funding this project:

the United States Congress.

Also, the reduction of flight time is of extreme importance because

the atmosphere of Pluto is expected to collaspe close to the year 2020. If

Intrepid should arrive after the collapse, the mission would be a failure; for

Pluto would be only a solid rock of ice.

i
SCIENTIFIC INSTRUMENTATION (SI)

Mission Science Objectives

Our desire to discover more about the Pluto/Charon system is the

driving reason this project was conceptualized. Due to the vast distances

involved, there is an extremely limited amount of knowledge pertaining to

the Plutoian system. Information that is presently available is subject to a

relatively high degree of error. Our Pluto/Charon mission will answer

questions regarding the system's age and origin, its classification as planet

and moon, internal dynamic interactions, and gather such planetary

characteristics as surface and atmospheric composition, magnetic field,

rotation rates, et cetera.

In addition to obtaining data from the Plutoian system, Intrepid will

observe Jupiter and examine the interplanetary space through which it

passes. The Jovian system will provide a gravity assist to the spacecraft.

During this time, Intrepid will point its far-scanning instruments towards

the Jupiter in order to gain more information about its complex planetary

system. The vast majority of the mission, however, will be spent in deep,

interplanetary space. Intrepid will gather data concerning the solar

winds, the interstellar particle medium, and the solar and interplanetary

magnetic fields. This information will help answer questions such as how

our solar system interacts with" the rest of the galaxy and where the

boundary of the solar system is. In addition, these investigations are

necessary for monitoring the calibration and performance of field

instruments.

£
Science Objectives at Pluto

Upon arrival at Pluto and Charon, the majority of science objectives

will be examined. The far-sensing instruments will begin collecting data

on the Plutoian system up to three months before the encounter date. These

instruments will study the transition between the interplanetary media and

solar wind/interstellar media, the interaction of Pluto and Charon's

planetary magnetic fields, the structure of the system's magnetospheres,

the ionosphere, plasma density profiles, rotation rates, and dynamic

interaction between Pluto and Charon.

As the encounter date nears, the remaining scientific experiments

will also be in operation. The atmospheric character will be closely

scrutinized. This includes any unusual features or haze, temperature and

pressure profiles, and composition. The surfaces of Pluto and Charon will

also be a major item of focus. The scientific instrumentation will carry out

a mapping of the surface geology, examine the polar icecaps, study the

color variations and albedo of the surface, determine cratering rates, and

composition of the atmosphere and surface. In addition, the imaging

equipment will take numerous photographic pictures of Pluto and Charon.

Accurate densities, masses, and radii of the planetary figures will be

determined.

All this information will be used to understand what type of

planetary system Pluto and Charon comprise. Whether Pluto is actually a

moon of Neptune which escaped, an oversized asteroid, or truly a planet

may be ascertained. Because it is the farthest planet from the Sun and has

never been examined by a spacecraft, there is very little known about it.

3
This mission will fill this void and perhaps provide more insight into the

formation of our solar system.

Limitations and Requirements of the Mission

Because of Pluto's extreme distance from the Sun, it is of prime

importance to arrive before Pluto's thin atmosphere collapses. This is

predicted to occur between 2020 and 2025. Should Intrepid arrive too late, it

will encounter only a large ball of ice. Thus in order to gain as much

valuable information as possible, it is of the utmost importance that the

spacecraft reach Pluto as soon as practical.

In order to beat the atmospheric collapse deadline, the Intrepid

spacecraft must be assembled and launched before 2005. Therefore, it must

be fully designed, built, and launched within twenty-five years. Plus, the

spacecraft must be assembled with reliable parts so it will survive the

fifteen year voyage.

The RFP has several explicit requirements that directly relate to the

SI subsystem. They are as follows:

1. design an unmanned science study to Pluto and Charon

2. mission science objectives must be stated and justified

3. design should stress reliability, simplicity, and low cost

4. spacecraft must be able to adapt to whatever environment it may

encounter

5. lifetime of spacecraft must include mission time and a safety

margin

6. nothing in the spacecraft's design should preclude it from

performing several missions

7. use existing hardware when possible in the design

8. use material and techniques available by 1999
In order to fulfill these inherent and explicit requirements, the

scientific hardware has been selected from previous deep space missions.

This decision to use existing hardware greatly reduces the development

costs involved, guarantees that they can survive the deep space

environment, and that they are indeed reliable. Fewer experiments have

been included than in previous missions in order to keep the spacecraft

design simple, low in mass and power requirements, and as inexpensive as

possible and while carrying out the science objectives.

The spacecraft missions from which the instrumentation originates

from should be: recent enough to have relatively current technology, not be

dependent upon much visible light, and have long component lifetimes.

The missions which fit these requirements best are the Voyager, Galileo,

and Mariner Mark H series. Although the Mariner Mark H series

(MMII) has not been launched, considerable research has been invested to

ensure that they are deep space worthy. Also, the MMII will be completed

well before the technology deadline of 1999.

The total component lifetimes is the largest stumbling block that

must be overcome with respect to science instrumentation. The scheduled

mission lifetimes of the Voyager, Galileo, and Mariner Mark H series are

approximately four, four and one half, and eight years respectively. Since

Intrepid's scheduled mission time is fifteen years, the scientific

instrumentation portion of the development costs will be spent upon

modifying the existing equipment so it will be able to surpass the lifetime

requirement. The mission lifetime can be a deceiving measure of the

actual lifetime performance. The Voyager's instruments, for example,

have been proven to be extremely reliable and are expected to exceed the
original mission lifetime many times. Therefore, the development costs

related to ensuring instrument lifetime should prove to be rather small.

Also, the science objectives may be slightly altered once the Hubble

Space Telescope examines the Plutoian system. It may provide knowledge

of large craters or surface anomalies that ought to be more closely

examined, a better approximation of the depth of the atmosphere, and more

accurate planetary eharacteristics. Once this data is analyzed, there may

be additional scientific equipment chosen in order to better study the

system. In addition, the existing equipment may be further modified to

incorporate additional objectives.

Science Instruments

The following science instruments have been selected:

Cosmic-Ray Detector System (CRS), from Voyager

Plasma Detector System (PLS), from Mariner Mark H (MMII)

Magnetometer (MAG), from MMII

Ultrastable Oscillator (USO), from Voyager

Solid-State Imaging (SSI), from MMII

Ultraviolet Spectrometer (UVS), from MMII

Figure 1.1 shows the scientific instrumentation layout. Appendix B

lists the masses and power requirements for these instruments.

The equipment selected from the MMII mission is not yet fully

developed. Mariner Mark H is also using available hardware designs and

upgrading them with more recent technology. The designs that they are

using, judging from the approximate masses and power demand

estimates, are from the Galileo spacecraft. (Draper, 10) Also, the MMII

has the longest lifetime of the missions considered. Therefore, when the
Solid-State Imaging

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C_0smicrRay Spectrometer

Plasma Detector System

Magnetometer

Ultras.tableOscillator

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Galilean instrument is better than that of the Voyager for the Pluto

mission, the MMII instrument with the more recent technology already

incorporated is the final selection. The Pioneer is not considered because of

the age of the technology incorporated in it.

The first three instruments listed above, the CRS, PLS, and MAG,

a_e particle or field scanning devices. These do not require accurate

pointing at a particular target body. The precise orientation of the

spacecraft needs to be incorporated with the instrument data for an

accurate record of the medium under study. These instruments are turned

on intermittently throughout the mission, and when Intrepid approaches

Pluto. The other three instruments are concerned with planetary science

and are turned on only when nearing a planetary target. The USO is a

supplement to the radio equipment; it is used in conjunction with the radio

instruments to conduct many radio science experiments. The USO does not

require any particular pointing or unobstructed view of space. The last two

instruments, the SSI and UVS, are best placed on a high precision pointing

platform because of their accurate pointing requirements.
The Cosmic-Ray Detector System

Cosmic-rays originate from planets and from stellar sources.

energies vary accordingly: (Flight Science Office, 4.1)

Their

Energy Origin

< 100 MeV/nucleus

1-100 MeV/nucleus

> 30 MeV/nucleus

Interstellar (from our Galaxy)

Nearby interstellar, or outer solar system

Jovian magnetosphere

Cosmic-ray composition past the giant planets is currently unknown. The

study of cosmic-rays in this region may provide insight on galaxial

composition and formation. Also, it will help scientists decide where the

boundary of the solar system is located.

The CRS has three separate particle telescopes for examining the

different energy ranges listed above. For the study of Pluto and Charon, the

telescope of the lowest energy range would be used. These individual

telescopes measure the charge, energies, and particle directions in their

respective energy range. In order to provide an unobstructed view for the

telescopes, the CRS has been mounted on the Science Scan Platform. Also

studied will be the cavities caused by the Plutoian system in the stellar

radiation. Because the CRS measures the charge composition of the

planetary magnetosphere, it provides a certain redundancy of data for

planetary magnetic fields. Therefore, not all the information concerning

the magnetic fields would be lost should the magnetometer become

damaged.

The reason for the selection of Voyager's CRS is, primarily, that it

fulfills the scientific objectives of cosmic-ray investigation. The Galileo
spacecraft instrumentation studied particles of less than 60 MeV. (Colin et

al., 6) Their primary purpose was to investigate the Jovian system, and

therefore did not have as great of a range as Intrepid's mission requires.

Thus, Voyager's design is the best choice.

The Plasma Detector System

Plasma is composed of mostly low-energy electrons and ions. The

plasma found in the solar system originates from stellar sources and from

the magnetospheres of planets. Because of the limited exploration of the

area, little is currently known about the composition of interplanetary

plasma beyond the giants.

The PLS measures this plasma and records its energy levels, ionic

composition, and velocities. Of particular interest for this mission is the

measurement of the magnetospheric plasma of Pluto and Charon. The

PLS will record the interactions between Pluto's and Charon's

magnetospheric plasmas with one another, and with interstellar and solar

winds. In addition, the PLS will identify the species of interstellar ions.

Also, the interaction of the interstellar and solar winds in the heliopause

will be studied.

Galileo's plasma experiment is clearly the better choice. It is able to

measure approximately ten times the energy ranges that Voyager could.

Its temporal resolution is twenty times faster, which is extremely

important for a flyby mission. Also, the PLS can produce a threedimensional vector velocity profile distribution of the plasma particles every

twenty seconds and identify species of interplanetary ions; which are tasks

the Voyager could not perform. (Colin et al., 133)
Magnetometer

Magnetic fields, studied

everywhere in the solar system.

by the magnetometer, are present

They are present on most planets, and

accompany the streams of charged particles which comprise the solar

wind. There is a great difference in strength between the magnetic fields of

planetary and stellar origin. Because of this variance, there are two sets of

sensors which are sensitive to the differing levels of intensity.

The MAG studies the magnetic fields from all origins: planetary,

solar, and interstellar. It measures the strength, fluctuations, and

structure of these fields. Of special interest is where these fields meet and

influence one another. The MAG also measures the heliopause accurately.

Because of the wide-spread presence of the magnetic fields, the

magnetometer is an important instrument for studying the large-scale

characteristics of the solar system. It provides much information on how

the solar system interacts with itself and the rest of the galaxy. For this

mission, the MAG will also determine if Pluto and Charon have magnetic

fields. If so, it will study the structure and characteristics of these fields.

The MAG's sensors are located on a separate boom of the spacecraft.

This location has been designed in order to minimize the detection of the

spacecraft's magnetic fields by the sensors.

Each magnetometer from the separate spacecraft has the same

purpose: to provide an in depth study of the magnetic fields. Therefore,

since the MMII incorporates the more current technology and has the

longest design lifetime, it is the one that Intrepid will use.

Ultrastable Oscillator (for radio science experiments)
Unlike the other instruments, the USO is only a supporter of a larger

experimental system. The telecommunication equipment is used to

preform a number of important experiments. The USO reduces the

transmitter frequency fluctuations to 1 to 4 x 10 -12. (Anderson et al., 228) It

drastically improves the accuracy of results from the radio science

experiments. The radio science is able to deduce the following information

by use of occultations and scintillations: temperature, pressure, and

density profiles of the upper atmospheres; electron density profiles and

irregularities in the ionosphere; magnetic field direction; gravity fields;

mean densities; bulk composition of the planets; and plasma density and

dynamics. Relativistic effects may be also be investigated by means of

comparing how signals of different frequencies from the spacecraft are

affected by the solar wind and corona. (Anderson et al., 224)

As Intrepid passes Pluto, the path will be such that Pluto is

positioned between Earth and the spacecraft. This configuration provides

the occultation of the radio signal3 required for many of the experiments

listed above.

In this case, the MMII used the same USO as did Voyager. (JPL

Mission Group) Because there is little available information about Galileo's

USO and the technology used for the Voyager is identical as for the MMII,

Intrepid will use the USO of Voyager.

Solid-State Imaa, in_

The SSI device is a combination telescope-camera which observes

and determines much about planetary atmosphere as well as its visual

physical characteristics. The SSI will be focused on the Plutoian system for

72 days before and after the encounter date. This prolonged viewing

/1
window, using long exposure times, will allow an in depth study of the

internal dynamic interactions of the Pluto/Charon system. Since Charon

rotates about Pluto once every 6.4 days, numerous rotations will be

observed. In addition, the SSI system will perform geographical mappings

of both planetary figures when closer to the system, using shorter exposure

lengths. Between these two modes of viewing, the following scientfic

objectives will be completed: locate the spin axes; record the dynamic

interaction of the system; obtain accurate measurements of the planetary

figure and size; study the surfaces' morphology, color, albedo, and surface

textures; measure the atmospheric energies via wave propagation modes;

help determine atmospheric radiative properties; search for possible

auroral interactions caused by magnetospheric interactions; and obtain

optical images of Pluto and Charon.

The Galileo design is a modification of the Voyager and Mariner 10

designs with the major upgrade being the use of charge-coupled devices

(CCDs). This along with other improvements yields an increase in

sensitivity of a factor of 100. (Hunten et al., 226) Also because of the

smearing problems, software was written to allow careful spacecraft

maneuvering while the camera is in use. (Fisk et al., 11) These

modifications will eliminate the problems Voyager had with image smear.

Even though there is very little light at Pluto, Intrepid should obtain

accurate optical data due to the CCD modification and that Intrepid is

passing within approximately 20,000 km of the system. Since this

instrument is extremely sensitive to movement, it is mounted on the high

precision pointing Science Scan Platform (SSP) in order to maximize

pointing accuracy. Another benefit of mounting the SSI on the SSP is to

guarantee an unobstructed view for the optical equipment.

/£
Ultraviolet Spectrometer

The Ultraviolet Spectrometer is used

atmospheric conditions. It is designed to

in the investigation of

study the atmospheric

composition, upper atmospheric atomic and molecular hydrogen, search

for ultraviolet emissions from the dark side of the planet to indicate any

auroral activity, and examine the cloud and haze structure. (Hunten et al.,

233-234) The UVS is a telescope-spectrometer with three detectors attached.

It will be located on the SSP so that the spacecraft need not perform any

special maneuvers for the pointing requirements of the UVS. It is not a

long range scanning device, thus it will begin scanning approximately 24

hours before the encounter date and will remain scanning for an additional

24 hours afterwards. The range of Pluto's atmosphere is unknown and

Charon's atmosphere, if it does exist, has not been confirmed. Therefore to

ensure that no features will be missed, the UVS will scan the entire

distance between the planets.

Although this distance is only 19,400 kin, the scanning fieldwill be

20,000 km long. This extra scanning range will reduce the uncertainty of

the Plutoian's system's measurements. To provide an accurate reading,

the UVS will remain fixed at an angle and allow the spacecraft drift to

move the fieldof vision. At Pluto Intrepid will be traveling 11.6 km/sec (see

MMPC subsystem), therefore, each scan of 20,000 km will last one half

hour. Using this method, almost one hundred scans will be recorded. The

Hubble Space Telescope will provide more data about the atmosphere before

launching. This would reduce the scanning length involved.

The reasons for selecting the Mariner Mark II over the Voyager

design are the improved wavelength ranges, longer expected lifetime,and

the more recent technology incorporated in the instrumentation.

/3
Anderson, Croft, Eshleman, Fjeldbo, Levy, and Tyler. "Radio Science

Investigations with Voyager." Space Science Reviews v 21 no 1,

1977.

Colin, Fanale, Franke, Hunter, Johnson, and Yeates. Galileo: Exploration

of Jupiter's System Washington, D.C.: National Aeronautics and

Space Science and Technical Information Branch, 1985.

Draper, Ronald F., JPL "The Mariner Mark H Program." AIAA-84-

0214, AIAA 22nd Aerospace Sciences Meeting Reno, Nevada,

December, 1984.

Fisk, L.A., L. Allen,and E.C. Stone, "Hearing Before the Subcommittee on

Science, Space, and Applications of the Committee on Science, Space,

and Technology, U.S. House of Representatives." One Hundred First

Congress, First Session. October 1989.

Hunter, Colin, and Hansen. "Atmospheric Science on the Galileo

Mission." Space Science Reviews v 44 no 3, 1986.

Personal Interview. Jet Propulsion Laboratory: Mission Design Group.

Pasadena, California, 17 April 1990.

Voyager Neptune/Interstellar Mission: Flight Science Office "Science and

Mission Systems Handbook, Revision D." Jet Propulsion Labs, 1987.
MISSION MANAGEMENT, PLANNING AND COSTING (MMPC)

Requirements of the Mission

The RFP for this mission has several design limitations that directly

affect the decisions and recommendations of the MMPC Subsystem. They

are as follows:

1. design an unmanned scientific study to Pluto and Charon

2. optimize cost and simplicity in design tradeoffs

3. launch time between 2000 and 2010

4. design should stress reliability, simplicity, and low cost

5. for cost estimates: assume four S/C built, three flight ready

and one for the integrated ground test system

The decision to develop a flyby mission, for example, was a direct

result of the consideration of the following RFP requirements: optimize cost

and simplicity in design tradeoffs; and stress reliability, simplicity and low

cost of the overall mission.

Type of Mission

The first step in the design of an exploratory spacecraft is to

determine the type of mission that will best satisfy the mission

requirements. The three types of missions which can be flown are: 1)

lander, 2) orbiter, and 3) flyby. Although each type of mission has both

advantages and disadvantages, it is the responsibility of the MMPC

Subsystem to determine which type will best satisfy the objectives and

requirements of the mission.

/S
The main advantage of a lander mission is the amount of time the

spacecraft is exposed to the planet. As a result of this large encounter time,

the largest (and most accurate) quantity of data is obtained. The drawbacks

of a lander mission, however, tend to outweigh the advantages. The large

spacecraft weight that accompanies a lander mission directly affects

several key mission requirements. Both the flight time and the change of

velocity (AV) are significantly increased as a result of the increased weight.

The mission driving factor (low cost), however, ultimately rules out the

lander as the spacecraft's mission type.

The orbiter mission is similar to the lander in that the spacecraft is

exposed to the planet for a great deal of time. Once again, however, the

drawbacks tend to outweigh the advantages. Although an orbiter mission

weighs less than a lander mission, the weight of the spacecraft (mainly due

to the amount of propellant needed to inject the spacecraft into a planetary

orbit) continues to affect the mission requirements. The mission driving

factor of maintaining low cost, therefore, rules out an orbiter mission as

well.

The flyby mission differs from the other mission types in that the

encounter time with the planet is greatly reduced. Although the amount

(and accuracy) of information obtained is less than the other mission types,

the advantages of using this type of mission tend to outweigh the

disadvantages. The small spacecraft weight that is characteristic of flyby

missions directly translates into a decrease in total AV, flight time, and

spacecraft cost. This reduction in spacecraft cost (along with the reduction

in AV and flight time) is the reason the flyby mission type was ultimately

selected over the lander and orbiter missions.

l0
Trajectory Determination

One of the many requirements the MMPC Subsystem is the

determination of a trajectory that will best fulfill the mission requirements.

This was best carried out using a program provided by SAIC (Scientific

Applications International Corporation) entitled MULIMP. Given certain

trajectory parameters (such as total mission time, arrival boundary

condition, and gravity assist bodies), the MULIMP program would find the

minimum AV trajectory. Comparing the outputs from several different

types of trajectories, one is able to decide upon the trajectory which satisfies

the mission requirements best.

The MMPC Subsystem considered numerous trajectories. The most

promising of these are: 1) Earth-Mars-Jupiter-Pluto (E-M-J-P), 2) EarthJupiter-Pluto (E-J-P), and 3) Earth-Earth-Jupiter-Pluto (E-E-J-P). The E-MJ-P trajectory was considered because of the Mars Initiative currently

under developedment. The E-J-P trajectory, as well as the E-E-J-P

trajectory, was considered because of Jupiter's large gravity assist

potential.

In the selection of the best trajectory, several parameters were placed

under consideration. The best combination of launch energy (C3), flight

time, nonlaunch _V (which determines the majority of the propellant

needed), and launch vehicle compatability (including its cost) will

determine the trajectory that will be used. Table 2.1 summarizes the three

trajectories under consideration (values were taken from MULIMP outputs

for each trajectory). From table 2.1, the trajectory that best fulfilled the

mission requirements (especially low cost) was the Earth-Earth-JupiterPluto trajectory.

t7
Table 2.1: Parameters for trajectory consideration

Trajectory

E-E-J-P

E-J-P

E-M-J-P

C3

Trajectory Parameters

Flight time Nonlaunch AV Launch Vehicle Compatible =

68.4

91.7

15.0 years

14.3 years

[.

27.6 13.2 years

1536 m/s

8353 m/s

5799 m/s

Titan IIID/Centaur

($130-140 million)

Titan IV/Centaur

($230 million)

Titan IV/Centaur

($230 million)

S . .

Least expenmve compatlble launch vehicle was selected

Trajectory Analysis

The Intrepid spacecraft will depart on its Earth-Earth-Jupiter-Pluto

trajectory in early February 2002. The parameters of the launch are given

in table 2.2 below ( values retrieved from the MULIMP program & the

Propulsion Subsystem).

Table 2.2: Launch parameters for the Intrepid spacecraft

Launch Parameter

i

Launch Energy

Launch AV

Launch (wet) Mass

Propellant at Launch

Value

i

68.4

5980 m/s

Launch Vehicle

Launch Vehicle Cost $130-140 Million

l

l S6kg

80O kg

Titan IIID/Centaur

18
The spacecraft is mounted in launch configuration on a Titan

IIID/Centaur launch vehicle (see Launch Vehicle Subsystem for launch

vehicle selection). The launch sequence begins with the firing of Titan's

solid rocket motors and first stage. Upon completion of the burn, the rocket

motors and first stage will fall off and the second stage will commence

firing. The second stage will fall off after the completion of its burn and the

Titan's payload fairing will separate, releasing the spacecraft and Centaur

upper stage.

The spacecraft and Centaur combination will coast for a length of

time for safety and proper trajectory insertion purposes. During this

coasting period, the Centaur will be spun in order to stabilize itself and the

spacecraft. After coasting for a sufficient period of time, the Centaur upper

stage is ignited. After completion of this burn the spacecraft will deploy its

booms. The booms are deployed before Centaur separation in order to

reduce the possibility of damage to the RTGs and science platform. The

Centaur upper stage will then be separated from the spacecraft with

pyrotechnics. The launch vehicle adapter will also be separated from the

spacecraft at this time (also with pyrotechnics). Finally the spacecraft is

stabilized and proper trajectory corrections are performed to place the

spacecraft on the desired trajectory. The launch sequence is summarized

in figure 2.1.

Approximately 2.9 years after launch, the spacecraft will return to

Earth and use it as a gravity assistbody. It willperform a AV maneuver of

1536 km/s at Earth to place iton the proper trajectorytoward Jupiter. The

AV maneuver along with the gravity assist will hurl the spacecraft toward

Jupiter by an additional 3650 m/s. Its closest approach to Earth will be

approximately 1.22 Earth radiior 1400 km from the surface. Nearly 1.45
Figure 2.1: Intrepid launch sequence

J

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years later,the spacecraR will encounter the Jovian system. However, the

spacecraft will rely entirely upon the gravity assist of Jupiter to aid in its

velocity since a AV maneuver is not needed. The spacecraf_ will gain an

additional 6700 m/s as a result of the assist. The closestapproach to Jupiter

is approximately 19.8 Jupiter radii or 1,400,000 km from the surface. This

distance is maintained in order to minimize the effect of Jupiter's intense

radiation and magnetosphere.

Approximately 10.7 years after the flyby of Jupiter (15 years since the

launch from Earth), the spacecrat_ will encounter the Plutoian system.

The spacecraft will be traveling approximately 11.6 km/s and it will be close

to 34 AU (5.1 billion kin) away from Earth. About a week before encounter
the spacecraft will direct itself to a point 20,000 km from Pluto's surface

(determined by the Scientific Instrumentation Subsystem to be a good

distance for studying the system). The thrusters will fire until the

spacecraft is on a direct path toward this point. The mission will end

approximately 72 days after the Pluto flyby, although extra fuel will allow

for additional burns if so desired. A possible post-mission of the spacecraft

may include the search for the heliopause (the point at which the Sun's

influence ends). The trajectory encounters are summarized in table 2.3.

Table 2.3: Summary of trajectory encounters

Earth

0

Earth Assist

2.9

Jupiter

4.35

Pluto

Time 1 (years) 15

AV 2 (kin/s) 5.978 1.536 0 0

AV 3 (kin/s) N/A 3.65 6.70 N/A

Dist. 4 (AU) 0 0 4.49 34.01

Dist. 5 (kin) N/A 1400 1,400,000 20,000

1 --time since launch

2 --AV performed by the spacecraft

3 --AV provided by the gravity assistbodies

4 --distance from the Earth

5 --closestapproach distance

Trajectory correction maneuvers (TCM) will be required to keep the

spacecraR on its trajectory. The TCMs will occur every 2-3 weeks on the

average and immediately before and after planetary encounters. As a
vehicle travels through space the trajectory is degraded (mainly due to

gravity gradients and solar flux). Correction against these degradations is

essential to ensure that the spacecraft will remain as close to its trajectory

as possible. This is especially important during' planetary flybys due to the

severe trajectory changes the vehicle endures.

Mission Costing

The process of examining cost allocations is a necessary ingredient

in the development of a new spacecraft. By breaking down major

subsystems into standard cost categories, MMPC Subsystems can

determine top-level cost estimates for these categories as well as for the

spacecraft as a whole.

The process of determining system cost estimates begins with the

calculation of direct and recurring labor hours (DLH and RLH

respectively). A summary of the equations involved are presented in

Appendix A2.1. Using these results and the conversion factors from

Appendix A2.2, the recurring (RC) and non-recurring (NRC) cost estimates

for the hardware-related categories can be calculated using the following

equations:

RC = RLH (labor hours to labor cost)(labor cost to total cost)

NRC = (DLH - RLH)(labor hours to labor cost)(labor cost to total cost)

These values are then used to determine a totalcost estimate (TC) for each

hardware-related category based on the SAI Planetary Program Cost Model

given in Appendix A2.3 (the X values used to determine Z were supplied by
the subsystems). The total cost estimates for the functional support-related

categories, on the other hand, are calculated using the following equation:

TC = DLH(labor hours to labor cost)(labor cost to total cost)

The total cost estimate can now be calculated by summing the

following category cost estimates:

* Development Project -Flight Hardware

* Development Project - Support Functions

• Flight Project

This total added to the cost of the launch vehicle used will result in the

spacecraft'stop-levelcost estimate.

Intrepid's top-levelcost estimate using the process described is $1032

million (based on the cost of building 4 spacecrafts). The category cost

breakdown is summarized in Appendix A2.4.
Appendix A2.1: Summary of Cost Model Algorithms (Koepke)

Development Proiect - Flight Hardware

Structure & Devices

DLH = 1.626 (N'M) 0.9046

RLH = 1.399 (N'M) 0.7445

Thermal Control. Cabling & Pyrotechnics

DLH = exp {4.2702 + 0.00608 N'M)

RLH = 3.731 (N'M) 0-6082

DLH = 56.1878 (N'M) 0-4166

RLH = 1.0 (N'M) 0.9011

Attitude _ Artigulation Control

DLH = 21.328 (N'M) 0-7230

RLH = 1.932 (N'M)

Telecommunications

DLH - 4.471 (N'M) 1.13o6

RLH = 1.626 (N'M) 1.1as5

Antennas

DLH = 6.093 (N'M) 1.1348

RLH = 3.339 (N'M)

Command & Data Handlin_

DLH = exp {4.2605 + 0.02414 N'M)

RLH = exp {2.8679 + 0.02726 N'M)
DLH = 65.300 (N'M) 0.3554

RLH = 7.88 (N'M) 0.7150

Landin_ Radar/Altimeter

DLH - 11.409 (N'M) 0.9579

RLH = 1.2227 (N'M) L2367

Line-Scan Ima_in_

DLH -- 10.069 (N'M) 1-2570

RLH = 1.989 (N'M) 1.4089

Particle & Field Instruments

DLH = 25.948 (N'M) 0.7215

RLH = 0.790 (N'M) 1-3976

Remote Sensin_ Instrumenta

DLH = 25.948 (N'M) 0.5990

RLH - 0.790 (N'M) 0.8393

Development Project - Support Functions

System Support & Ground Equipment

DLH = 0.36172 (ZDLHhardware) 0.9815

Launch + 30 Days Operations & Ground Softwar_

DLH = 0.09808 (_DLHhardware)

Ima_ng Data Development

DLH = 0.00124 (Pixels-Per-Line) 1.629

Science Data Development

DLH = 27.836 (non-imaging science mass) 0.3389
Pro Lrram Management/MA&E

DLH = 0.10097 (ZDLHall categories) 0"9670

Flight Proiect

Flight Ouerations

DLH -- (ZDLHhardware/3100)0.6(10.7 MD + 27.0 ED)

Data Analysis

DLH = 0.425 (DLH Flight Operations)

N

M

DLH

RLH

MD

ED

- Number of spacecrafts

- Mass in kg

- Direct labor hours in 1000 hours

- Recurring labor hours in 1000 hours

- Mission duration in months

Encounter duration in months
Appendix A2.2: Labor/Cost Conversion Factors (Koepke)

Cost Category Labor Hours to Labor Cost 1

(Feb'90 dollars/manhour)

DevelopmentProject

Structure & Devices 22.04

Thermal Control, 21.64

Cabling, & Pyrotechnics

Propulsion 22.23

Att. & Artic. Control 22.42

Telecommunications 21.07

Antennas 21.01

Command & Data Handling 20.41

RTG Power 20.06

Landing Radar/Altimeter 21.26

Line-Scan Imaging 22.29

Particle & Field Instruments 22.40

Remote Sensing Instruments 22.46

System Support & Ground Eq. 22.25

Launch + 30 Days Ops & Gr. S/W 22.59

Image Data Development 24.17

Science Data Development 26.91

Program Management]MA&E 24.40

Flight Project

Flight Operations 22.02

Data Analysis 22.02

Labor Cost to Total Cost

3.303

3.317

3.616

3.347

3.352

3.466

3.163

3.177

3.158

3.604

3.395

3.286

3.076

3.214

3.130

3.987

2.685

3.247

3.425

1 - Feb'90 dollars= 2.109(FY77 dollars) (US Dept of Commerce 462 & US Dept ofCommerce 6)
Appendix A2.3: SAI Planetary Cost Model (Koepke)

Inheritance Class Categories

• Class One: Off-the-Shelf/BlockBuy

• Class Two: Exact Repeat of Subsystem

• Class Three: Minor Modifications of Subsystem

• Class Four: Major Modifications of Subsystem

• Class Five: New Subsystem

Cost Reduction Algorithm bv Inheritance Classes

Let X1 -

X2-

X3-

X4-

X 5 -

Percent of Subsystem Off-the-Shelf

Percent ofSubsystem Exact Repeat

Percent of Subsystem Minor Modifications

Percent of Subsystem Major Modifications

Percent of Subsystem New Design

Thus Xl + X2 + X3 + X4 + X4 + X5 = 100% of Subsystem Mass

NRC

RC

TC

Z

= Non-recurring cost estimate (without inheritance)

-- Recurring cost estimate

- Total cost estimate (including inheritance effects)

- Percent cost reduction

If Z = 1.0Xl + 0.8X2 + 0.25X3 + 0.05X4 + 0.0X5

Then TC = (100% - Z) NRC + RC

7_
Appendix A2.4: Intrepid cost estimates

Cost Category Cost (in millions of dollars)

Structure & Devices

Thermal Control, Cabling & Pyro.

Propulsion

Attitude & Articulation Control

Telecommunications

Antennas

Command & Data Handling

RTG Power

Landing Radar/Altimeter

Line-Scan Imaging

Particle & Field Instruments

Remote Sensing Instruments

System Support & Ground Equip.

Launch + 30 Days Op & Ground S/W

Image Data Development

Science Data Development

Program Management/MA&E

Flight Operations

Data Analysis

$22.85

$10.13

$24.53

$32.55

$35.68

$11.18

$56.40

$25.92

$o.82

$142.21

$33.36

$0.72

$215.29

$73.46

$5.03

$9.41

$70.52

$83.92

$37.62

Subtotal

Launch Vehicle

$891.6

$I,I0.0

Top-level cost estimate $1031.6
l CES

Koepke, Andrew. AAE 241 - Aerospace Vehicle Design Class Notes.

University of Illinois, Spring 1990.

MULIMP. Electrical Engineering/CSO lab cite.

U.S. Dept. of Commerce/Bureau of the Census. "Statistical Abstract of the

United States." 109th Edition, 1989, p. 462.

U.S. Dept. of Commerce/Bureau of Economic Analysis. "Survey of Current

Business." Vol 70 No 3, 1990, p. 8-6.
POWER AND PROPULSION b-XTBSYb-_I_'VI(PPS)

Propulsion Subsystem

The propulsion subsystem must provide the spacecraft with the

ability to make AV as well as course correction maneuvers and attitude

control adjustments. This subsystem may be divided further into three

categories: propellant and pressurant, propulsion feed system, and

thrusters. The breakdown of masses for the entire propulsion subsystem

can be found in Appendix B.

Prooellant and Pressurant

For the spacecraft,two types of chemical propellants are considered.

These are monopropellants and bipropellants. Solid fuels are immediately

ruled out because of their nonexistent stop-restart capabilities which are

essential for a Plutoian mission. Furthermore, cryogenic fuels are

neglected because of their poor storage qualitiesover long durations, again

a requirement for a mission to Pluto.

A monopropellant system will be used on the spacecraft because of

the advantages it has over a bipropellant system. First, monopropellants

require less complicated propellant feed systems. This results in mass

savings, and consequently cost savings, due to the reduced tankage and

valving required. Also, many monopropellants are storable for long

durations; bipropellants are not. Bipropellants do however have a distinct

advantage over monopropellants in specific impulse. This advantage will

reduce the amount of propellant required for a given mission. However, a
driving factor in choosing a propulsion system is simplicity and reliability.

Overall, a monopropellant system better fits the RFP requirements than a

bipropellant system for a long duration mission.

The spacecraft fuel selected is monopropellant hydrazine (N2H2),

with a specific impulse of 225 seconds. Not only does hydrazine provide the

lowest cost propulsion system (Koepke), it is simple, reliable, and has been

used on a number of previous spacecraft. Additionally, hydrazine has been

shown to be space storable for extended periods of time (greater than 12

years) and has stop-restart capabilities (Koepke). Furthermore, thrust

levels from as low as 0.2 N to moderate levels of approximately 2500 N have

been demonstrated using hydrazine (Koepke).

Despite the advantages of using hydrazine for the spacecraft's

propellant, disadvantages also exist. Drawbacks to hydrazine include its

moderate plume contamination of the spacecraft and its toxicity. The

problem of plume contamination is easily remedied by the use of shielding

and strategic placement of sensitive instruments. The dangers of handling

hydrazine have been greatly reduced because of the familiarity gained with

the fuel from previous use in other spacecraft.

Estimates based on trajectory analysis performed by the MMPC

(Mission Management, Planning and Costing) subsection indicate that the

total AV propellant required for the mission is 677 kg, while the total

trajectory correction and attitude correction propellant is 65 kg. A

contingency of 58 kg of extra propellant was included in order to account for

2.0% unusable fuel (14.84 kg) and an error margin for propellant

consumption. Thus the resulting total propellant mass is 800 kg.

The pressurant for the propulsion system is used to force the

propellant into the propellant feed lines. Two types of pressurants were
considered for the spacecraft: helium and gaseous nitrogen. Although

each is reliable,cheap, and proven, helium was selected due to its mass

savings: 2.24 kg as opposed to 15.63 kg for GN2. Additionally its lower

liquifyingtemperature is better because fewer heaters are required.

Provulsion Feed System

The prime requirement for the propulsion feed system is to supply

the thrusters with propellant. This system can be further reduced into:

tankage, valving, filtering,and tubing. A schematic of the propulsion feed

system is given in figure 3.1.

_ Reralator

PillValve _.-._ /Valw

-.-----.---_ P_ief Vadve

(4) 1_ N Tl_nmten

Servi_ Valve,

I

fl2q 0.8 N Ttu'_t_n

Pigure 3.1. Propulsion System Configuration
Both the main trajectoryand attitude correctionfuel willbe contained

within one main tank. This provides for a simpler and lighter system

although redundancy is sacrificed. However, research found no evidence of

spacecraft failures due to faulty propellant feed systems thus the design is

considered adequate.

A regulated pressurant feed system was selected over a blowdown

system. For a blowdown system, the tank inletpressure decreases over the

lifetime of the mission resulting in a reduction of thruster performance. A

reduction in performance decreases the specific imptLlse of the system

(figure 3.2).

3OO 3OO

r ..=.... "-'-'- _ 2oo

2oo

150 _ _J 150

_. ,,-

5O 5O

50 250 450 10 20

INLET P_ESSURE <PSlAI

t

J

30 40

THRUST ILBI

50

Figure 3.2. Inlet Pressure versus Specific Impulse (TRW)

This loss in performance would require the use of more propellant which

would negate any mass savings achieved by using less valving and tubing.

Two spherical, titanium tanks were designed for propellant and

pressurant storage. A spherical design was chosen because it provides the

greatest volume to surface area ratio,it can withstand large stresses using

a minimal amount of material, aids in the stabilityof the spacecraft, has

been flight tested, and may be an off-the-shelfitem which is currently

available.
The propellant tank is approximately 1.2 meters in diameter and is

designed for a pressure of 250 psi (1.7236 MPa) with a safety factor of 1.5.

The design pressure was chosen based on desired thruster performance

plots (figure 3.2). The tank can hold 800 kg of hydrazine and will use a

spherical bladder system. This will help keep the center of mass of the tank

from moving as fuel is drained. A titanium alloy,Ti-6AI4V, is chosen for

the tank material because of its resistance to damage from the hydrazine

fuel,its moderate density (4400 kg/m3), itshigh yield strength (850 MPa),

and its proven reliability(Ashby, 10). Although research has not indicated

that a tank as previously defined already exists,there is confidence that

such a tank may already be available.

The pressurant tank has a diameter of 0.50 meters and is designed

for a pressure of 3000 psi (2.068 MPa) again with a safety factorof 1.5. This

tank will hold 2.24 kg of helium pressurant. The same titanium alloy as

used for the propellant tank will be used here resulting in a tank mass of

16.15 kg. Again, this type of tank may already be a'railable.

Aluminum alloys, particularly the Al 7000 series, were considered

for the pressurant tank material giving a mass of approximately 18.39 kg,

however material costs would be reduced by $95.54 (1988 dollars)over the

titanium tank (Ashby, 10). This dollar savings was not considered

sufficientin order to account for the much larger specificcost expected at

launch.

The valves, filter,and tubing mass required for the system was

estimated based on the Voyager spacecraft (Mangano). All the material

used here will easily be found as off-the-shelfhardware which has been

previously flighttested.
Thrusters

Both the main propulsion system and attitude control system require

the selection of thrusters. While the attitude system uses low-thrust

thrusters in order to make minor attitude adjustments, the primary

propulsion system will use larger thrusters to provide the necessary AV

maneuvers.

For the main thrusters, four 100 N thrusters were selected providing

a total of 400 N. A four thruster configuration provides redundancy

whereas one thruster does not. The MRE-50 monopropellant thruster

manufactured by TRW (TRW) is an off-the-shelf,flight proven thruster

capable of meeting the design requirements. Care was taken sizing the

thrusters in order not to perturb the spacecraR structure during thrusting

(maximum accelerations of" approximately 0.06g are expected).

Additionally, the duration of the AV burns must be within the design limits

of the thrusters. The AV maneuver requiring the longest duration burn

will occur at the spacecraft's earth gravity assist. This burn will last

approximately 3736 seconds which can be accomplished by pulsing the

thruster nine and one-thirds times with each pulse lasting 400 seconds.

Since the selectionof thrusters is not final,other types may be substituted to

better fitthe requirements. However, the basic design should remain the

same.

For the attitude correction maneuvers, the AAC (Attitude and

Articulation Control) subsystem has determined that twelve 0.2 N thrusters

similar to those used on Voyager, willbe adequate. Again this is an off-theshelf item which has been flight tested and proven reliable on earlier

spacecraft.
Further Comments

An additional feature to the propulsion subsystem design includes

autonomous control. Since the time for one way communication with the

spacecraft at Pluto will be on the order of four hours, pre-programmed AV

maneuvers based on the spacecraft's location (determined from attitude

control) must be included. Furthermore, the spacecraft will be passing

Pluto at a relative speed of approximately 11.6 km/sec so manual trajectory

corrections will be impractical.

Placement of the thrusters is also of importance. Four 0.2 N

thrusters will be placed on the outer rim of the bus ninety degrees apart.

Orientation has been determined by the AAC subsystem. The remaining

eight 0.2 N thrusters and four 100 N thrusters are located on four separate

pods underneath the spacecraft (opposite the high gain antenna). Each pod

has one 100 N thruster and two 0.2 N thrusters. These pods will be

arranged in a square pattern 1.5 meters apart. This location and

orientation was chosen in order to reduce plume contamination to sensitive

instruments, provide control in the event of a thruster failure,and avoid

interference with the adapter structure. Shielding will also be used to help

reduce plume contamination.

Since the overall design of the propulsion subsystem is general and

simple, nothing should preclude it from performing other missions.

However, this design may be better suited for long duration missions in

order to help justifythe use of a monopropellant. Propellant requirements

for other missions will dictate whether this spacecraft design would be

feasible.

7
Any changes to the propulsion subsystem would affect the structure

and AAC subsystems primarily. Specifically, changing the type of

propellant, thruster system, or tankage would have the largest impact.

Care should be taken ifsuch changes are necessory.

The lifetime of the spacecraft is indirectly affected by the propulsion

subsystem. If a major failure occurs, the spacecraft would be unable to

correctly orient itselfand would essentially become "paralyzed". Of key

importance is the thruster lifetime. It is particularly dependant on the

frequency and duration of burns so these must be kept to a minimum to

ensure an appropriate spacecraft lifetime.

Power Subsystem

The power subsystem may be divided into the following categories:

power generation, energy storage, and power conditioning. A schematic of

the power subsystem is shown in figure 3.3 while a breakdown of power

requirements for each subsystem is given in table 3.1. Furthermore,

Appendix B provides a listingof the masses for the power subsystem.

I P_er I

Marmrmnmt

l

Source _m'_ Control and

Cantr_ Di_ibut_on

Modular _tmt l_t_r power Control Umt

RTG'e Power D4otributton Umt

Ftuml

Power l

Pr_Neol_

I 2.4 kl_

Figure 3.3. Schematic ofPower Sub_let4ma
Subs_,stem Cruise Pre-Fire Maneuver Science Comm.

Telecom 10 10 10 65 65

CCC 15 1,5 2O 25 25

TCS 20 20 20 20 20

AACS

SI

9O 9O 90 90 9O

19 19 19 56 56

TOTALS 154 154 159 256 256

Table 3.1. Breakdown of Power Requirements [watts]

Power Generation

Power generation is achieved using a modular radioisotope

thermoelectric generator (MRTG) design. RTG's provide the only feasible

source of power for the distances and trip times that the spacecraft will

encounter on a mission to Pluto. In addition to providing essential

redundancy, a modular design also offers a weight savings over a nonmodular design because of the abilityto tailorthe power source to the power

requirements. Furthermore, if the power requirements are changed, only

adding or subtracting RTG "slices"is required for meeting the new power

demands. This is especially important if the spacecraft is to have multimission capability. Using RTG's do however have serious drawbacks.

These include thermocouple degradation and public concern over the

RTG's radioactivity.

For the spacecraft's mission to Pluto, it is determined that the

maximum required power is 256 W (required at Pluto). In order to account

L
for the degradation of the thermocouples and the half-life of the plutonium238 fuel, an extra 75 W will be required at launch (degradation of 5 W/yr

(Fisk, 6) over a trip time of 15 years). Thus, the total power required at

launch is 357 W.

Specifically, the MRTG design here is comprised of fifteen individual

"slices" (figure 3.4) each producing 24 W of power at 28 V with a specific

power of 10.57 W/kg (Schock, 341). This produces a total output of 360 W and

a MRTG mass of 34 kg. It should be noted that although the power

contingency margin appears to be narrow, each subsystem's requirements

includes an additional contingency.

Figure 3.4. Modular RTG Design (Schock, 338)

Concerns on the safety of the RTG's during launch is also of prime

importance in the design. However, since there is extensive testing done on

RTG containment in the event of a launch failure,problems are not likelyto

occur.

Energy Storage

The selection of an energy storage

consideration. Two devices are considered:

controller.

system is an important

batteries and a discharge

Although both provide excellent energy storage qualities,itwas
decided that an energy storage device would not be incorporated into the

spacecraft design. Because the spacecraft willgenerate more power than is

required during all phases of the mission, any power shortage may draw

on the contingency power. Furthermore, the power management system

will reduce power to those components which are considered less essential

in the event that more power is required than that available from the

contingency. This allows a weight savings by not including extra

components and keeps the design simple.

Power Conditioning

Power conditioning entails the following: power source control,

power control and distribution,power processing, and power management.

All the components used in this subsection are off-the-shelfitems that are

taken from the Voyager design (Mangano).

Power source control is achieved through the use of a regulated

power system. The primary component for this system is a shunt

regulator. This design was chosen because it provides a constant voltage

supply and has been proven on previous spacecraft. A disadvantage to a

regulated system is that it requires more mass than an unregulated

system. However, this added mass can be justified because a regulated

system is a proven system.

The power control and distribution component consists of a power

control unit, a power distribution unit, and fuses. All three prevent power

overloads which may damage the spacecraft's electronics. Additionally,

the power control and distribution units control the power to various

subsystems during maximum power periods.
Power processing consists of inverters for changing the RTG's DC

generated power to AC power. The primary unit used here is a 2.4 kHz

inverter.

The power management system has control over the entire power

system. Power cycling,which helps reduce the chance of power shortages

and extends the lifetime of the various electroniccomponents, is controlled

through this system. Also included in the power management system are

parallel connections and redundancy which reduce the chance of single

point failures. Lastly, autonomous control of the power management

system will be provided by the command subsystem in case of a loss of

communications with the ground stations, or when communications are

too long to permit appropriate control.

Further Comments

The lifetime of the entire spacecraft is dependant primarily on the

lifetime of the RTG's thermocouples. As these degrade, the spacecraft's

power supply is reduced. The power management system is designed to

cycle the available power to the various subsystems in order to alleviatethe

problem of reduced power and extend the useful lifeof the spacecraR.

Changing the power subsystem primarily affects the scientific

instrumentation and communications subsystems. Of particular concern

are changes to the RTG's. For this design, tailoringof the power supply is

easily accomplished by using the modular RTG's. This helps the

spacecraft adapt to various mission designs.
Launch Vehicle Subsystem

The launch vehicle has the responsibility of giving the spacecraft

enough energy to begin its mission on the proper trajectory. The following

are requirements for choosing the spacecraft'slaunch vehicle:

1) itmust provide an adequate C3 (launch energy) for the given

spacecraft launch mass

2) itmust adequately fitthe spacecrat_ and any upper

stages within the launch vehicle

3) itmust be as inexpensive as possible

4) itmust be reliable.

Various launch vehicles were considered. The two that best fitthe above

requirements are the Titan IIID / Centaur and the Titan IV / Centaur. The

corresponding payload dimensions and cost are given in table 3.2 while

launch energies are given in figure 3.5. Keeping the previous requirements

in mind, the Titan IIID/Centaur was selected as the launch vehicle (figure

3.6).

Launch Vehicle

Titan IIID/Centaur

Titan IV/Centaur

Payload Dimensions

3.65 m diam, 10.7 m length

4.33 m diam, 8.93 m length

Cost (1989 dollars)

$130-140 million

$240 million

Table 3.2. Launch Vehicle Information +

+ The Titan IIID/Centaur payload dimensions are in fact Titan Ill Commercial

specifications(Gizinski,4). The Titan IIID/Centaurinformationwas not available.The

TitanIIICommercial dimensions(fora dedicatedlaunch)areexpectedtobe smallerthan

thosefortheTitanIIID.TitanIV information(AviationWeek, 163).
\, i

i Pq.Al_ TARY _E RF'Ci_AA_a_

_, "_ _'. _. -- ........

Figure 3.5. Titan IIID/Centaur and Titan IV/Centaur Launch Energies

versus Launch Mass (Atkins, _ and JPL Mission Group)

Figure 3.6. Titan IIID/Centaur*

Both launch vehicles provide the spacecraft,(mass at launch = 1386

kg) with the required launch energy of 68.4 km2/sec 2. Additionally, the

payload fairings for both adequately fit the spacecraft in its launch

configuration with the Centaur upper stage. Other features similar for

both launch vehicles include their outstanding success rate (94% reliability)

and launch site (Cape Canaveral) (Gunn). The cost of the Titan IV /

Centaur however was nearly twice that of the Titan IIID / Centaur. This

factor is the deciding reason for choosing the Titan TTID / Centaur.

* Figure 3.6 is actually that of a Titan IIIE/Centaur (Heacock, 214). The Titan

IIID/Centaur looks similar to the Titan IIIE/Centaur.

ORIGINAL PI_GZ IS

OF POOR QUALITY
Further Comments

The choice of using an expendable over a reusable launch vehicle is

driven by the requirements of low cost and simplicity. Any increase in the

spacecraR's mass, however, willnot allow the use of the Titan IIID due to

launch energy constraints. Some other type of launch vehicle,such as the

Titan IV ! Centaur must be used.

Any changes to the launch vehicle selectionshould not greatly affect

the design of the spacecraft as long as payload fairingand launch load

requirements are met. However, ifa more powerful launch vehicle (higher

C3 capacity)is selected,changes in the planetary trajectorymay be allowed

perhaps reducing the flighttime of the mission.
Appendix A.3 - Equations

Propulsion Subsystem

• Equation@ for propellant mass and volume reouired:

mpropl AVJ = (mdryS/C + 1/3mIAVTCM) (exp(AVJ/Isp) - 1)

mprop ]AVE = (mdryS/C + 2/3m IAVTCM + mprop I AVJ) (exp(AVE/Isp) - 1)

mproptotal ffimprop IAVJ + mpropl AVE + m lAVTCM

Volume of propellant = mass propellant / density of propellant

For a spherical tank: Vtank = 4/3*PI*r 3

• Equation for vressurant mass and volume reauired:

repress ffi(atomic weight of pressurant) * (mambpx of moles of pressurant)

# of moles = (PtankVtank)/(RpressTtank)

• Provellant and _ressurant tank sizin_ eauations

mspherical tank = (safety factor)*(2PI)*Ptankr3(tank material density/yield stress)

• Thruster sizin_

acceleration = ms/c/thrus t

tburn = (DmIspg)/thrust
Atkins, Kenneth Lee. On the Electric Prouulsion/Mission System: _nera_.

Constraints. and Characteristics Surfaces. Doctorate Thesis.

University of Illinois, Aeronautical/Astronautical Engineering

Department. August 1974.

Ashby, Michael F. and David R.H. Jones. En_neerin_ Materials 2: An

Introduction to Microstructure. Processing. and D¢@ign. New York:

Permon Press, 1988.

Fisk, L.A., L. Allen, and E.C. Stone, "Hearing Before the Subcommittee on

Science, Space, and Applications of the Committee on Science, Space,

and Technology, U.S. House of Representatives." One Hundred First

Congress, First Session. October 1989.

Gizinski, S.J. III and D.B. Herrington. "Titan III: Commercial Access to

Space." AIAA-88-3476, 1988.

Gunn, C.R. "AIAA: Space Transportation Technical Committee Status of

NASA Mixed Fleet Programs." Presentation given to TRW Space &

Technology Group, 5 May 1989.

Heacock, Raymond, JPL. "The Voyager Spacecraft." Proceedings of the

Institution of Mechanical En_neers v 194, 1980.

Koepke, Andrew. AAE 241 - Aerospace Vechicle Design Class Notes.

University of Illinois, Spring 1990.

Mangano, M. Division 35 Mass Estimation Reference Revision - Interoffice

Memorandum # 35261/80/01, Jet Propulsion Laboratory, 24

November 1980.

MRE-50 MonoDrooellant Thruster. Brochure: TRW Space & Technology

Group, 1989.

NASA Space Systems Technoloaw Model. NASA Office of Aeronautics and

Space Technology. NASA TM 88174, June 1985.

Personal Interview. Jet Propulsion Laboratory: Mission Design Group.

Pasadena, California, 17 April 1990.

Schock, A. Modular Isotopic Thermoelectric Generator. Germantown,

Maryland: Fairchild Space and Electronics Company.

"U.S. Launch Vehicle Specifications." Aviation Week and Space

v 132 no 12, 19 March 1990.
STRU_ SUBSYSTEM (SS)

The structure subsystem (SS) is further divided into three smaller

subsections: structural design (SDS), thermal control (TCS), and materials

(MS). The driving requirements of the SS are low cost (i.e. low weight),

simplicity, and reliability.

Structural Design Subsection

The SDS consists of the layout of all components, launch vehicle

compatibility, launch load requirements, and on-orbit assembly. Adapting

previously flight tested designs offers greater reliability and less redesign

work. Pioneer 10 & 11 provided invaluable data to the Voyager program in

eliminating concerns with the asteroid belt hazard, defining radiation

environments at Jupiter, and demonstrating that a spacecraft (S/C) can

survive in space for extended trip times. Similarly Pioneer, Voyager,

Galileo, and the Mariner Mark H (MMII) program give Intrepid

invaluable data. Therefore to reduce cost, the Intrepid design uses existing

technology whenever possible.

Figure 4.1 shows the primary subsystem components. Numerous

iterations were performed to satisfy the structural requirements of the

complex payload against competing requirements of the entire system.
Science

Imaging camera

Magnetometer

Plasma Detector

UV-Spectrometer

Ultra Stable Oscillator

Cosmic Ray Detector

ParticleDetector

Power/Propulsion

RTG

Propulsion Tank

Various Thrusters

Communication

High Gain Antenna

Low Gain Antenna

Computer

Attitude Control

Platform

Computer

Various Thrusters

Figure 4.1 Subsystem Necessities

Item

All Items

Magnetometer

RTG

Science

Problem

Keep Cost Low

Interference

Radiation/Thermal

Inertial

Provide Unobstructed View

Inertial

Solution

Off Shelf Hardware

Use Existing Technology

Maximize Distance From S/C Bus

Mount On Boom, Shield

Balance Science

Mount On Boom

Balance RTG

Figure 4.2 Structural Design Considerations

Figure 4.2 shows the structural design considerations of various

components. During the structural design phase, these considerations

were taken into account, as well as the placement of the center of mass

(CM) and the component inertial contributions to the spacecraft (S/C). The
CM placement and inertial contributions are also very important

considerations for the attitude/articulation control subsystem (AACS).

These values were calculated and optimized using the INERT program

(INERT, AAE 241).

Intrepid Spacecraft Design

The configuration of the Intrepid S/C (without thermal blankets) is

shown in Figures 4.3 - 4.5. The basic structural component of the S/C is a

ten bay metallic bus which houses the electronics. A trade study for the

MMII program was performed examining different bus structure

configurations and the number of electronic bays. The study concluded that

the bay approach is more mass efficient(Draper, 7). Also, for the bay

packaging approach, all required testing fixtures, procedures, and

experience presently exist, reducing risk of developing and cost. Thus, the

b_y approach was chosen.

Inside the bays, the electronics are packaged on flat plate

(sandwiched aluminum honeycomb) sub-chassis which, when installed

vertically, become integral structural elements providing strength to the

spacecraft (Heacock, 217). This packaging allows for more desirable CM

properties. Another consideration is the use of a rectangular group of bays

or a toroid design. Rectangular groups have the advantage of having a

small end so thermal shielding on one face of the mission can be quite

small (Draper, 7). However, the MMII program found that the use of

rectangular bays leads to a larger structure than when using a toroid. A

reduced mass plus experience with a toroidal bus design are the driving

factors for selecting the toriod design.

5O
Figure 4.3 Intrepid Spacecraft Flight Configuration (x-z plane)

Figure 4.4 Intrepid Spacecral_ Flight Configuration (x-y plane)

I---- . _ _. "---'-t

T

tql,_

4_

[" i

_--'_ o k _ / '

it "_o0_ d

ORIGINAL PAGE IS

OF POOR QUALITY

5_
Figure 4.5 Intrepid Spacecraft Flight Configuartion (y-z plane).

P'_"__N

./

\ /

• " / i

1

U:

The last bus design consideration is the number of bays required.

Galileo used eight bays, but had several packages of electronics outside the

bus. Due to the trip duration (15 years), a design which houses electronics

on the outside of the bus is not prefered. Therefore, a ten bay bus design is

selected,similar to the Voyager design.

Power/propulsion subsystem (PPS) selected the use of modular

Radioisotope Thermal Generators (RTG)s. The RTG placement is governed

by their gamma ray and neutron radiation and the need to balance the

mass of the science platform. Science instrumentation subsystem (SI)

selected various instruments which are mounted on a scan platform,

....... _ PPI,"#" |S

52
providing mass balancing of the RTGs about the S/C CM. Furthermore, the

field of view for the sensors is maximized. To reduce the effects of the RTG

radiation, the science instruments are mounted on the science boom 180°

from the RTG boom. An added benefit of the 180° design is that the main

structure of the S/C will help shield the science from exposure to radiation.

Another device reducing the radiation exposure is a radiation shield on the

RTG boom. These techniques help reduce the length of the booms needed to

adequately separate the RTGs from the science.

Reducing the length of the booms also helps alleviate vibrational

problems. The science and RTG booms are 2.1 and 3.5 meters long,

respectively. The use of dampers will also reduce vibration problems.

Any electrical current flowing within a S/C can produce a magnetic

field sufficient to distort measurements of weak interplanetary fields.

Similar to the Voyager S/C, the low field magnetometers are mounted on a

10 meter Astro-Mast boom in order to minimize interference from the S/C's

magnetic fields. The magnetometer is directed 900 from the other booms.

Additional strategies for reducing magnetic interference include proper

selection of materials and strategic placement of all necessary electrical

apparatus.

The remaining components: high gain antenna (HGA), low gain

antenna (LGA), and propulsion tank are positioned based on previous S/C

and CM placement requirements.

Launch Vehicle Compatibility

Once the total weight of the S/C was fixed (see Appendix B for

calculations) the PPS selected the Titan III-D/Centaur launch vehicle for

launch.
r

i

k

Figure 4.7 Intrepid Spacecraft in Launch Configuration

OPdC.'.L_]ALPA_E ]_

OF POOR QU/k.LrrY

Figure 4.7 shows the Intrepid S/C in its launch configuration,

excluding thermal blankets. The Titan III-D has a payload bay diameter of

3.65 m, a payload length of 10.7 m. The upper stage Centaur is 5.9 m long

with a resulting maximum S/C length of 4.8 m (Gizinski IH, 4). As can be

seen in Figure 4.7 the Intrepid S/C (in launch configuration) fits within the

allotted dimensions.

Launch Vehicle Survivability

In the launch configuration, the primary S/C structure (bus and

propellant tank) are connected to an upper stage assembly (USA) adapter.

The S/C and supporting adapter will experience maximum accelerations of

4g's during launch, producing a maximum of 52,960 N of force. Since the

adapter's graphite/epoxy truss elements can withstand 59,270.4 N of axial

compression, the S/C meets launch load requirements (see appendix

A.4.2). During launch, the adapter structure will be used to support the

large mass of the RTG and science platform booms. This is required to
prevent damage from launch phase acceleration, vibration, shock, and

acoustic environments (Heacock, 216).

On-Orbit ass¢rably

The mission's on-orbit assembly consists of deploying the RTG,

Science, and magnetometer booms. The RTG and science booms are

deployed by actuators, the adapter is separated by pyrotechnics, and the

magnetometer boom is deployed.

The magnetometer and boom are stored in a canister that is .23 m in

diameter and .66 m long. The boom is a triangular truss made of fiberglass

longitudinal members held in place by fiberglass triangles. These are

stiffened with tensioned, collapsible diagonal filaments. The boom is

stowed by twisting the entire structure so that the diagonal filaments

interlace and the triangles are nearly in contact with each other. This puts

a considerable elastic force on the assembly. When the canister is opened

by a pyrotechnical device, the boom expands at a controlled speed to prevent

the boom from popping out with destructive violence (NASA, 20-21).

Thermal Control Subsection

The TCS is responsible for the regulation of temperature on board the

S/C. The temperature in deep space is -273.33 °C (Genovese, 2), while the

operating temperature of most instruments is between -20 °C and 40 °C. In

addition,TCS policygenerally requires that components be operable at 25 °C

above and below the predicted temperature extremes (Braun, 5). This

difference creates a criticaldesign problem. Some considerations for the

TCS are gradual decreasing temperatures away from the sun, frigid

periods (i.e.Earth shadow, Jupiter shadow), engine heat, and RTG heat.
Another important design consideration is the degradation of instrument

and S/C optics and thermal control surfaces. Therefore, highly sensitive

surfaces are oriented so they receive minimal ultraviolet radiation and

thruster plume influence (Braun, 3). Also all lower thrusters will have

plume shields to protect sensitive surfaces from damage.

Table 4.1 Method of Achieving Thermal Control

Passive

Active

Heatin_

Multi Layer Insulation

Surface Paints (Black)

Placement of Heat Generatin_ Equipment

ElectricHeaters

Radio-isotopeHeaters

Coolin_

T

Spring-actuated Louver

Surface Paints (White)

Placement of Heat Generating Equipment

Shunt Radiators

Table 4.1 lists the devices that will be used for temperature control in

the Intrepid S/C. Passive means are employed whenever possible to reduce

cost. Passive devices include bimetallic spring-actuated louvers, multilayered insulation (MLI) blankets made of aluminized Mylar or Kapton,

various surface paints, and placement of heat generating equipment.

Active means include electrical heaters, radio-isotope heaters, and

temperature sensors. Existing hardware elements from previous S/C have

been used to provide a low-cost, low-risk, thermal control design.

The TCS was further subdivided into two smaller subsections: bus

and science thermal control (BSTC), and the propulsion unit thermal

control (PTC).
Bus & Science Thermal Control

The BSTC subsection controls the bus, science boom and the antenna

thermal enviornrnent. The antenna is painted white to control its

temperature in near-Earth (1 AU) solar environment. Additionally, an

electricallylossy paint will be used to prevent electrostaticcharging of the

surface (Heacock, 2i7). The bus thermal control will be achieved using

several strategically placed supplemental heaters, louvers to release

internal heat, and MLI blankets. The science boom thermal control uses

supplemental heaters and an MLI blanket taking into account operating

temperature differences from instrument to instrument. The MLI blanket

will be constructed with the aluminized side toward space. Thereby the

solarinput to the S/C increases by six foldin surface absorptivity (Braun, 4).

The bus and science temperature will be controlled to be -10 °C to 30 °C,

within the required operating temperature.

Propulsion Unit Thermal Control

The PTC system can be divided into two distinctzones: the interior

hydrazine propulsion bay and the strut mounted outboard thruster

modules. A minimum requirement of 8.33 °C is applied to all propellant

delivery items, propellant tankage, and thruster valves to provide an

adequate margin and therefore preventing hydrazine freezing. Maximum

temperature limits were established to ensure structural and functional

integrity during the mission. These limits are 65 °C (fluid),148.88 °C

(thrusters),to 37.77 °C (propellanttank) (Genovese, 3). The propulsion bay

temperature will be controlled by multiple heater/thermostat circuits and

thermal isolation by three MLIs (Genovese, 2). The thruster modules are

thermally controlled according to specificsections. The fluid distribution

57
and control elements such as propellant feed lines and valves will be

conductively insulated, actively controlled with heaters and thermostats,

and radiatively isolated with MLI blankets, while the thrust chamber will

be equipped with heaters and sensors (Genovese, 3).

Materials Subsection

The MS selects the materials to be used for different subsystems. The

MS is also in charge of micrometeorite protection. Table 4.2 mentions

different environments which will dictate the choice of material for the S/C.

These environments include: radiation; cyclic temperature changes; high

vacuum; vacuum outgassing; contamination; and particle

debris/micrometeorites (Pope, 760).

Table 4.2 Environmental Considerations

Hostile Environment_

Radiation

Cyclic Temperature Changes

High Vacuum

Vacuum Outgassing

Contamination

Micrometeorites

Desirable material properties to satisfy these demanding conditions

include: high specific modulus and strength; good radiation resistance;

high vibrational damping characteristics; low thermal expansion; and low
density (Pope, 760). Low density is important as it is directly related to

weight and therefore cost. Other considerations include low development

and overall cost, previous space worthiness, and space environment

lifetime. All the materials examined displayed good overall material

properties. All materials were selected on the basis of cost and that they

were flight proven. Table 4.3 gives the materials considered versus

requirements.

Table 4.3 Materials Considered

Aluminum

Beryllium

Titanium

Composites

-Graphite/epoxy

-Carbon-carbon

-Metal Matrix

-New fiber resin

-Sol-gel

-triphasic

Svace Tested Overall Cost

yes

yes

|0w

hi_h/medium

medium

Develooment Cost

N/A

N/A

_'es N/A

_,es low N/A

little medium medium

little medium

no

no

no

high

hil[h

high

medium

high

high

high

Graphite/epoxy was selected as our primary material.

Graphite/epoxy is flight proven and requires little developmental cost.

Composite materials provide high payoff in systems which require high

thermostructural stability because of their low coefficient of thermal

59
expansion and

consisting of

reinforcement

high thermal conductivity ("NASA", 2-12). Composites

high-temperature thermoplastics with graphite-fiber

provide light weight, good dimensional stability, and

excellent resistance to corrosion, chemicals, and wear (Dreger, 52).

Graphite/epoxy will be used for the bus bays, the adapter trusses, the

science boom, the RTG boom, and all supporting trusses.

The RTGs exterior will use beryllium. Beryllium offers a great

combination of weight, strength, and thermal conductivity. Beryllium is

one of the lightest structural metals and is used satellite structures. The

drawback of using beryllium - toxicity, special design needs, handling

requirements, and machining techniques - make it expensive to use. The

unique RTG requirements of high temperature, radiation, and severe space

environment dictates the use of this more expensive material.

The propulsion system's high temperature, and high stress require a

special material. The propulsion system will be made mainly of 6A14V

titanium alloy. Titanium is used in bo*_h the tank and thruster strut

network. Titanium is space tested, has low development cost and was

selected for the propulsion system for its low thermal conductivity and high

strength. The drawback of using titanium include slightly higher costs

and higher weight density.

Other materials used include: fiberglass (magnetometer boom),

aluminized Mylar (MLI thermal blankets), honeycomb aluminum material

(antenna), A120 3 coating (protective coating), and lossy surface paints

(thermal control).

Fiberglass is selected for use in the magnetometer boom because of

its unique launch configuration and on-orbit assembly. Fiberglass was also

used for the Voyager magnetometer boom thus reducing development cost
Appendix A.4.2

Launch loads and axial loads carried by adapter truss.

Maximun acceleration felt during launch (from propulsion

subsystem): 4g

(where g--9.8 m/s 2)

USA truss is able to carry 13,230 Ibf in axial compression (Stang, 1)

(13,230 Ibf )*(4.48 Newtons / 1 Ibf) ,, 52970.4 Newtons

Force on one truss:

trusses)

S/C launch mass., 1351 kg

acceleration = 4g = 4*(9.8 m/s 2)

Ftr * (# of trusses) = 52959.2 Newtons

Since the force the USA can carry is greater than the force on one

truss, the truss will be able to withstand launch loads.

59270.4 N > 52959.2 N

Ftr - (mass of S/C) * (acceleration) / (# of

Note: during launch the adapter withstands most of the forces acting

on the S/C. Reducing the stress on the actual S/C structure.
and ensuring reliability. Aluminized Mylar is the established material for

MLI blankets and has been used on several deep space missions.

Micrometeorite Protection

An additional concern of the MS is micrometeorite protection system

(MPS). Though guarding against a large asteroid would be futile,

protection from penetration of small micrometeorites is very important.

Previous S/C have proven flightworthy for extended flighttimes (Voyager

13 + years), therefore our MPS uses existing flight proven, reliable,

technology. The Intrepid S/C will use micrometeorite protection shields

while aluminized Mylar MLI adds additional protection. A similar system

is now being used on Galileo.

Further Comments: Conclusion and Recommendations

Possible problem areas include the need to do more precise inertial

calculations. The boom vibrational problem must be examined more in

depth, and the problem of the electronics vibrating inside the bus must be

studied.

The main concern of this proposal is to keep cost low and the design

simple, yet reliable. This was achieved through the use of off-the-shelf

hardware whenever possible. The SDS used previous S/C as their starting

point in design. TCS used established passive means of thermal control

where ever possible to reduce cost. MS avoided new untested composites

which would bring development cost up. The simple, reliable, and low cost

Intrepid is practical and has a long line of successful design predecessors.
References

Braun, C.E., and Budd, D.W., RCA Aerospace and Defense, Astro-Space

Division. "The Thermal Control Challenge of the Mars Observer

Spacecraft." AIAA-87-1501, AIAA 22nd Thermophysics Conference,

Honolulu, Hawaii, June 1987.

Draper, Ronald F., JPL "The Mariner Mark H Program." AIAA-84-

0214, AIAA 22nd Aerospace Sciences Meeting, Reno, Nevada,

December 1984.

Dreger, Donald R. "Plastics for Those Hot Jobs." Machine Design v 56 no

17, 26 July 1984.

Genovese, J.E. and Godward, J.L., United Technologies Corporation,

Hamilton Standard Division. "Thermal Design of a Large Spacecraft

Propulsion System." AIAA 87-1863, AIAA 23rd Joint Propulsion

Conference, San Diego, California, June 1987.

Gizinski, S.J. III and D.B. Herrington. "Titan III: Commercial Access to

Space." AIAA-88-3476, 1988.

Heacock, Raymond, JPL. "The Voyager Spacecraft." Proceedings of the

Institution of Mechanical En_neers v 194, 1980.

INERT Computer Program. Electrical Engineering Lab, CSO: UIUC,

1990.

Mangano, M. Division 35 Mass Estimation Reference Revision - Interoffice

Memorandum # 35261/80/01, Jet Propulsion Laboratory, 24

November 1980.

NASA, Scientific and Technical Information Office ed. Voyager to J_jpitcr

and Saturn. Washington, D.C.: NASA, 1977.

"NASA Space Systems Technology Model." June 1985.

Pope, Edward and Mackenzie, John, UCLA, Dept. of Materials Science &

Engineering. "Novel Composite Materials for Space Structures and

Systems." 32nd Internal SAMPE Symposium, April 1987.

Stang, D.A., Martin Marietta. "Advanced Composites for Magellan

Spacecraft." AIAA-88-3031, AIAA 24th Joint Propulsion

Conference, Boston, Massachusetts, July 1988.
COMMAND,CONTROL, AND CO--CATIONS (CCC)

Communication Subsystem

The communication subsystem must fulfill the following functions:

transmit information from the S/C to Earth (telemetry); receive information

from mission management (command); and provide information on the

spacecraft position, velocity, and radio propagation medium (tracking)

(Yuen, 1-3).

The communication subsystem can be subdivided into the Radio Frequency

Subsystem and the Antenna Subsystem. A detailed list of masses and

power requirements appears in Appendix B.

Radio FreQuency Subsystem

Intrepid's Radio Frequency Subsystem (RFS) will employ X-band

(8.414 GHz downlink and 7.161 GHz uplink) to fulfill all comm-mication

and radio science functions. Although the Voyager and Galileo

predecessors successfully communicated using X-band as well as S-band,

the foremost reason for utilizing only X-band is the decreased cost and

simplicity of calibrating one less frequency for operation through the Deep

Space Network (DSN) (Draper, 15). In addition to the higher recurring DSN

costs of other bands, the testing and fabrication of these bands is more

expensive and may require increased power for transmission (JPL Mission

Group).

Fundamental to the sole employment of X-band is the X-band

transponder. The usage of this transponder as opposed to an S-band uplink

possesses the following advantages: decreased ionization effects that
accompany the utilization of a higher frequency will result in increased

rate and range measurements; charged particle interference is reduced

thereby facilitating command of the S/C when approaching Solar

conjunctions; in the two-way coherent mode, doppler tracking accuracy

and bit-error rate are improved. This last advantage is due to a higher SNR

on uplink and virtually no continuation of the uplink frequency to the

downlink frequency. Fewer sources of Radio Frequency Interference (RFI)

exist on X-band than S-band and channel assignments are 50 MHz wide for

X-band as opposed to 10 MHz wide for S-band. The above trade studies were

conducted for the MMII program (Draper, 15).

The other major component of Intrepid's RFS is the X-band Solid

State Amplifier (SSA). Due to its greater simplicity, the SSA was selected

instead of the Traveling Wave Tube Amplifier as used on Voyager and

Galileo. In high power mode, the output of Intrepid's SSA will be

approximately 40 watts and have a mass of 5.5 kg. Details of mass and

power requirements for the RFS are located in Appendix B.

Antenna Subsystem

Intrepid's 1.5m High Gain Antenna (HGA) and 0.Sin Low Gain

Antenna (LGA) will both operate on X-band during uplink and downlink.

The technology for the redesign of a LGA to be compatible with the RFS is

based on research conducted during the MMII Program. This LGA will

operate near Earth until the Sun-S/C-Earth angle falls below 10 degrees.

Then the LGA will function mainly during emergencies and when

reorientation of the S/C is required (Draper, 21).
Technology for the HGA is also borrowed from the MMII Program.

The driving reason for utilizing the LGA and HGA technology from MMII

is to reduce cost. Although Intrepid will travel 14 AU farther than any

MMII mission, the 1.Sm HGA was selected. This decision was based on its

low weight and launch vehicle compatibility. Even though greater power is

generally required to transmit signals as the area of the HGA is reduced,

this problem can be offset with the employment of the X-band transponder

and the SSA in the RFS (Draper, 20).

Table 5.1 lists selected performance parameters for the Intrepid S/C

radio systems (Andrew, 61&67) (Anderson et al, 228):

TABLE 5.1

Performance Parameters for Intrepid Radio System

Transmitting Parameters:

Transmitting Frequency

Transmitter Powers

X-band Low Power

X-band High Power

Transmitting Antenna Gain

0.5 Meter Parabola

1.5 Meter Parabola

8.414 GHz

10 watts

13 watts

35dB

41 dB

Receiving Parameters:

Receiving Frequency

Receiving Antenna Gain

0.5 Meter Parabola

1.5 Meter Parabola

7.161 GHz

34dB

40dB

(rf_
Further Comments

Intrepid's antenna subsystem may need to be redesigned. Either a

larger HGA or possibly the addition of one or two Mid-Gain Antennas

(MGA) may be necessary to overcome the vast distances involved in order to

successfully transmit and receive signals from Pluto.

Another problem area concerns the life-timeexpectancy. Intrepid's

communication subsystem relies heavily on technology developed for the

MMII Program. Intrepid's trip time is 15 years - 7 years longer than any

MMII mission. However any developmental design costs to extend

Intrepid°s life-timeshould be minimal. In addition, life-time expectancies

are generally underestimated as witnessed by the success of the Voyager

mission. Voyager was designed for a life-timeof 4 years - itis now entering

its13th year oftravel.

Command and Control Subsystem

The Command and Control Subsystem (CCS) must monitor the

health of the entire S/C, ensure semi-autonomy and provide stabilityand

control for the SIC. This subsystem is further divided into the following

classifications: Command and Data Subsystem; Computer Command

Subsystem; Flight Data Subsystem; Data Storage Subsystem; Attitude and

Articulation Control Subsystem; microprocessors; data rates; and data

memory.
Comm_nd and Data Subsystem

Two single module computers will comprise Intrepid's Command

and Data Subsystem (CDS). Each computer will be programmed identically

to provide redundancy and will contain the Computer Command

Subsystem (CCS), the Flight Data (FDS) and Data Storage Subsystems

(DSS), and the :_,ttitude and Articulation Control Subsystem (AACS).

Detailed information regarding the AACS is outlined in the Attitude and

Articulation Control section.

Although Intrepid's computer system is modelled after Voyager's,

the following differences exist between them: Intrepid employs two

computers instead of Voyager's six; and each Intrepid computer contains

the CCS, FDS, and AACS whereas Voyager employed a separate computer

for each subsystem (Adamski, 2-3). Because of Intrepid's greater memory

capabilities (1 Gigabyte each), only two computers are necessary.

Comvuter Command Subsystem

The CCS is the principal controller of the S/C: The CCS relays

instructions to the AACS which controls the S/C attitude, sensors, gyros,

scan action, and thrusters; instructs the FDS to record and compress

images; and directs the DSS to record and/or playback the DTR. However,

the CCS's foremost responsibilities are the following: collect science data

and execute instructions commanded from the ground to operate the S/C;

and respond to any problems that materialize with other subsystems

(Kohlhase, 40-1).

This second CCS function is executed through Expert Systems (ES).

The ES contains software routines called Fault Protection Algorithms

[1
(FPAs). Approximately 20 percent of the CCS memory is comprised of

FPAs to allow the S/C to be semi-autonomous. Therefore the S/C can

respond to problems as they materialize (Kohlhase, 40). The

implementation and reliable performance of ES is crucial to Intrepid's

mission since such vast distances are involved; roundtrip light time may be

as high as eight hours. Since ground stations will communicate very

infrequently with Intrepid, once a month during cruise mode and then

more frequently as Pluto draws near, further S/C autonomy is mandatory.

Flight Data Subsystem

Pluto science and engineering telemetry data are gathered and

formatted in the FDS for transmission to Earth. The science data will be

downlinked at 300 bps whereas the engineering telemetry data will be

transmitted at approximately 40 bps. This engineering data is essentially a

status report of the health of the different S/C subsystems. Other

information provided are the S/C attitude and scan platform position

(Kohlhase, 42).

The FDS will utilize the Reed-Soloman (RS) encoding process. This

process, developed during the Voyager II mission, reduces overhead to

about 20% by adding only 1200 bits to every 3600 bits of raw science data sent

back to Earth. In addition, the number of bit errors is reduced from 5 in

100,000 to only 1 in 1 million (Kohlhase, 126-7).

The FDS includes the software routine of Image Data Compression

(IDC) developed for Voyager II. By counting only the difference between the

brightness of successive pixel grey levels rather than the entire picture,

IDC can reduce by at least 60% the number of bits that characterize each

.4

t _t'_-- I
image. This process reduces the time needed to transmit a complete image

from Pluto to Earth (Kohlhase, 126).

The FDS operates in a "dual processor mode" whereby the principal

mode formats the general science data and the secondary mode compresses

the images. By functioning in parallel and executing different functions,

the computer memory is effectively doubled. This allows for higher

computing tasks (Kohlhase, 42).

The scientific instrumentation and the digital tape recorder are also

controlled by the FDS. Such variables as filter choices, exposure times, and

imaging shutter modes are determined by the FDS (Kohlhase, 42).

Data Storage Subsystem

The Digital Tape Recorder (DTR) is the integral component of the

DSS. Since Intrepid°s return rate is a low 300 bps, the DTR must contain

sufficient memory to store all of Pluto's images. Memory storage of 880

Megabytes of memory will be required to fulfill Intrepid°s Pluto science

imaging objectives. At Pluto encounter, three speeds of the DTR are in use:

1500 bps (record only); 300 bps (playback only); and 100 bps (record and

playback).

Microorocessors

As modelled after the CRAF S/C to be launched in 1995, Intrepid's

CDS will incorporate a SA3300 16-bit solid state microprocessor with less

than 1 MIPS and a RAM of 128 or 256 kilobytes (Bowlin, 14). According to a

study conducted at Energy's Sandia National Laboratory in 1989,

approximately 5 years will be required to design, build, and test this space-
qualified microprocessor and its accompanying hardware (Bowlin, 15).

Final selection of a microprocessor will not be mandatory until 1997 at

design lock-in. This will ensure that the computer hardware can be

developed, tested, and integrated into the S/C subsystems to create a fully

operational S/C (Bowlin, 16).

The employment of a 16-bit microprocessor will allow for increased

on-board processing of mission data thereby reducing the amount of

necessary space-to-ground communications (Bowlin, 18). Improved onboard computer processing speed and memory allows for the use of a highlevel programming language further reducing software development costs.

Because of this, the S/C computers will be programmed in C. In addition,

less rewriting of software will be required upon upgrade of the S/C

computer subsystem (Bowlin, 19).

To fulfili our low cost mission objective, data rates will be severely

reduced as compared to VoyagerII. Whereas Voyager H transmitted 115.2

kbps in high data rate mode and 266 bps in low data rate mode (Kohlhase,

124), Intrepid will transmit from Pluto at a low rate of 300 bps. Low rate

return requires less power for transmission and therefore reduces costs

(JPL Mission Group).

Upon destination, all science information will be gathered and

temporarily stored for later low rate return. Picture data will then be

returned over a period of two to three months. At closest approach, 1.5 to 2

passes will be made through the DSN per day. Because of this, base

stations on Earth at Goldstone, Madrid, and Canberra will be able to

V_
provide 8 hours each of coverage. When only one pass is made through the

DSN per day, one base will be utilized at a time, greatly reducing recurring

DSN costs (JPL Mission Group).

As stated previously, reduced data rates will decrease costs and

require less power for transmission. Estimated power requirements

include 65 watts of raw power for a 13 watt transmission at 8.414 GHz.

More than 110 watts of raw power for a 22 watt transmission at 8.4 GHz

were required during the Voyager mission (Edelson et al, 921). These

values assume a 20 percent efficient radio frequency transmitter.

Data Memory_

The DTR will comprise the Data Memory Subsystem (DMS) and will

utilize technology from the MMII Program (Draper, 20). Due to Intrepid's

low data rate return, it is imperative that its tape recorder have ample

storage capabilities. A memory of 880 Megabytes will be required to fulfill

our Pluto science objectives (JPL Mission Group). Memory capabilites of

each computer will consist of 1 Gigabyte.

Further Comments

The success of Intrepid's Command and Control Subsystem relies

heavily on the SA3300 16-bit microprocessor. If the design, testing, and

development of this microprocessor is delayed and therefore unavailable at

design lock-in, an alternate microprocessor must be selected and further

redesign of the computer system must be initiated.
Adamski, Terrence, JPL. "Command and Control of the Voyager

Spacecraft." AIAA-87-0501, AIAA 25th Aerospace Sciences Meeting

Reno, Nevada, January 1987.

Anderson, Croft, Eshleman, Fjeldbo, Levy, and Tyler. "Radio Science

Investigations with Voyager." Space Science Reviews v 21 no 1,

1977.

Andrew Corporation: Andrew Catalog 34: Systems Planning ProduCt

Specifications Services. Orland Park, IL: Andrew Corporation,

1988.

Bowlin, Samuel, United States General Accounting Office. "NASA

Advanced Spacecraft Computers Report." Washington D.C.:

1989.

GAO,

Draper, Ronald F., J-PL "The Mariner Mark H Program." AIAA-84-

0214, AIAA 22nd Aerospace Sciences Meeting Reno, Nevada,

December 1984.

Edelson, R., B. Madsen, E. Davis,and G. Garrison. "Voyager

Telecommunications: The Brodcast from Jupiter." Science v 204 no

4396, June 1979.

Heer, Ewald and Henry Lure, eds. _[._chine Intelligence and Autonomy

for Aerospace Systems. Progress in Astronautics and Aeronautics,

Volume 115. washington, D.C.: AIAA, 1988.

Kohlhase, Charles, ed. The Voyager Neptune Travel G_id¢. California:

NASA/JPL, 1989.

Personal Interview. Jet Propulsion Laboratory: Mission Design Group.

Pasadena, California, 17 April 1990.

Yuen, Joseph ed. Dee_ S_ace Telecommunication_ Systems Engineering.

California: NASA/JPL, 1982.
ATTITUDE AND ARTICUI_TION CONTROL SUBSYSTEM (AACS)

Introduction

The driving design factors for the AACS, in addition to the

requirements stated in the request for proposal, are the strict pointing

accuracy and stability requirements needed for the spacecraft to

communicate back to Earth and to perform the mission's science

experiments. The AACS provides spacecraft orientation from the time of

separation from the launch vehicle through the end of the mission. The

subsystem must be able to control the spacecraft during all mission modes

and adapt to changing mission requirements. And possibly the most

important factor in the design of the subsystem is that it must be able to

perform autonomously during all mission modes. A system to meet these

requirements has been developed using standard technology and off-theshelf hardware.

Method of Stabilization

When selecting the method of stabilization to be used for Intrepid, the

mission requirements had to be examined closely. The nature of the

mission calls for stringent pointing accuracy and stability while the request

for the mission calls for low cost, simplicity and reliability. These factors

have been taken into consideration, and have led to the selection of a 3-axis

stabilization technique.

The 3-axis technique has been selected for its excellent pointing

accuracy, reliability and versatility (Koepke, 36). The pointing accuracy of a
three-axis spacecraft is only limited by the sensors used. The quality of

attitude sensors has increased substantially in the last few years, thus

improving the accuracy of the 3-axis spacecraft. This improvement will be

spelled out later when the attitude components are described. 3-axis

systems have been flight tested for long duration missions (Voyager), thus

the reliability of such a system is proven. The lifetime of the system is a

factor of the on-board propellent and the lifetime of the components. There

is going to be enough fuel on-board to carry out the mission plus

contingency, and the components have a lifetime exceeding the length of

the Intrepid mission, 15 years. The 3-axis method is much less complex,

less costly, and more reliable than the dual-spin method. Complexity and

the cost seem to go hand in hand. The more complex the system, the more

costly and less reliable it is. Dual-spin is not as reliable as a 3-axis system

because of the substantial increase of moving parts which are subject to

wear over a very long mission such as Intrepid's. The versatility of the

spacecraft is a very important aspect for a long duration mission. The 3-

axis system is more adaptable to changing mission requirements than the

spin-stabilized system. Furthermore, the 3-axis system has greater

maneuverability than the spin-stabilized spacecraft which gives it better

versatility when performing science procedures. The spin-stabilized

spacecraft has imaging limited to the line of scan. This makes it more

difficult to scan objects, which is a major objective of the mission. Also, the

configuration of the spin stabilized spacecraft is constrained by its

geometry, where as the 3-axis spacecraft is not. This gives greater

versatility to the design and layout of the spacecraft.

not

The disadvantage of heavy, costly hardware for 3-axis spacecraft is

as apparent. The new generation of attitude hardware for 3-axis

7_
systems is coming down significantly in mass, power and cost. This makes

the method of stabilization even more attractive.

High Precision Scan Platform (HPSP)

The nature of the mission requires very stringent pointing accuracy

and stability. The mission also entails that the spacecraft have the

capability to meet demanding and changing mission requirements. One of

these is the stability required to do imaging in the low level of light that will

be encountered at Pluto. The low level of light forces the camera to use

longer exposure times for imaging. The longer exposure time dictates a

need for increased stability to eliminate image smear. The increased

pointing accuracy and stability of the spacecraft call for a better method of

achieving these requirements. This increased pointing accuracy and

stability can be achieved using new technology from the Mariner Mark H

program, the high precision scan platform (Bell, 807).

The HPSP provides a rigid, 2 degree of freedom, momentum

compensated platform for precision inertial and celestial sensors and

science instruments with high pointing requirements (Bell, 807). The

momentum compensation in the platform allows the activity of the scan

platform to become dynamically isolated from the spacecraft bus. Thus, the

bus is not disturbed by scan platform motion. This is advantageous because

it allows scan activity without an impact on fuel consumption, thus

creating a savings on mass and cost. The HPSP gives greater pointing

accuracy and stability due to its rigid mount and decoupling. This

---7_
eliminates many sources of error in instruments that are evident on

existing platforms.

The HPSP can slew about two axes, azimuth and elevation relative to

the boom (see figure 1). This motion is accomplished by using an internally

redundant microstep actuator for each axis of motion (Bell, 807). Each

actuator has a reaction wheel that is automatically activated to compensate

for the torque created by the platform motion. This allows the dynamics of

the platform to be isolated from the bus. For attitude determination, the

HPSP contains an inertial reference unit and high accuracy star tracker.

The attitude is determined on the HPSP and then is transferred to the

spacecraft bus. The drawback of the HPSP is seen because it creates a

complex coordinate transformation to provide the spacecraft bus attitude

(Draper, 12). Thus, a sophisticated algorithm must developed to provide

quick relay of data from the platform to the microprocessor to the attitude

actuators. The total HPSP pointing accuracy is estimated to be 0.76

mrad/axis. This a vast improvement over that of the Voyager scan platform

accuracy of 2.26 mrad/axis (Bell, 808). A layout of the HPSP with the

attitude and science instruments is given in figure 6.1.

Inertial Reference Unit (IRU)

The inertial reference unit provides 3-axis rotation rate data for the

spacecraft in all control modes. The conventional IRU's used on past

spacecraft have employed mechanical gyroscopes. These, due to their

mechanical nature, are subject to drift error and must be updated

frequently. However, the length of the mission calls for maximum

autonomy and accuracy. So a new type of IRU is going to be used which can

7f
give this autonomy and accuracy with a mass and cost savings. This IRU is

called the Fiber Optic Rotation Sensor (FOILS) (Draper, 801).

FOILS is capable of a long life with high reliability (Stokes, 162). The

reliability of the sensor is a function of the light source used. This

uncertainty has been eliminated by using a semiconductor light source that

requires no thermoelectric cooler (Stokes, 163). FORS contains no high

power components gnd also uses standard, proven technology for the space

environment.

Azimuth

-91

_ I-_-.25--1l..

.16 1 2

1 Camera

2 ASTROS II

3 HPSP

4 UV Spec.

5 Cos. Ray. Sys.

6 FORS

all measures in meters

A Transverse

I

I

---. 9 ..--

4

•i.--_-. 41------_-

.25

.2

v

5

Figure 6.1 HPSP with attitude and science instruments

7f
FORS has many advantages over other types of gyros available,

namely mechanical and laser gyros (Stokes, 161). One advantage over the

mechanical gyros is that FORS does not have any moving parts. This allows

FORS to be active much longer without having to update for drift errors as

with the mechanical gyros. This translates directly into a cost savings. Also

with no moving parts, FORS is not subject to mechanical failure due to

wear on parts, thus creating a longer life span. A second advantage over

mechanical gyros is that there is no gravity sensitivity. This makes FORS

absent from additional error while the spacecraft is in a high gravity

region, such as near Earth or Jupiter. There are many advantages over the

ring laser gyro which are very important for the Intrepid mission. A high

voltage supply and mechanical dither are not needed in the FORS design.

And in addition, the FORS is available at a much lighter weight. All of

these advantages add up to a longer lifetime and lower cost for the FORS

inertial reference unit. Table 6.1 shows the drift, weight, and power of the

three types of sensors.

Rate Sensor

Residual Drift

Rate

(deg/hr)

2 * 10 4

Rate noise

(deg/hr)

1 * 10-5

Power(W) Mass (kg)

FORS 10 10

DRIRU II 3 * 10-3 1 * 10 -5 22 11

Laser Gyro 7 * 10-3 4 * 10 4 18 18

Table 6.1 Comparison of rate sensors (Draper, 14)

-7o
FORS has been created with the same interface as the NASA

standard gyro DRIRU II and is to be placed on the HPSP. Therefore the cost

of creating a new interface is eliminated (Bell, 803). Since FORS is not flight

tested, although it is expected to be flown on the Mariner Mark H missions,

Intrepid can revert back to the standard gyro easily if any problems surface

with the FORS unit.

FORS also measures body acceleration during thruster burns to

sense body rates. This eliminates the need for accelerometers (Stokes, 161).

The unit is 3-axis internally redundant, provides a long life, low mass and

solid state inertial rate and position sensor, and is being used as the inertial

reference unit for the Intrepid mission.

Star Sensor

The main requirements for attitude components has been reliability

over a long life span and redundancy. The only star sensor that will be

available that can meet these requirements while also giving optimum

performance, low mass and power consumption, all for a low recurring

cost is the Advanced Star and Target Reference Optical Sensor (ASTROS

II) star tracker (Draper, 15).

ASTROS II has increased value in many ways to the existing types of

trackers available. The main advantage is that it is fully internally

redundant. None of the other available star trackers have this feature,

which is very important for the length of our mission. Other advantages are

its high accuracy, low recurring cost, low mass and low power
consumption. It is expected to be flight tested on the Mariner Mark II

missions thus becoming off-the-shelf hardware.

ASTROS II utilizes a Charge-Coupled Device (CCD) which enables

closed loop pointing of small bodies such as Pluto or asteroids. This gives

greater flexibility if there are changing mission requirements because it

allows for autonomous science gathering operations (Bell, 803). The CCD

also gives ASTROS II the ability to have a bright particle pass in front of it

and not lose a star lock. This is very important for autonomy and cost since

it cuts down on the number of maneuvers to regain star lock.

Sun Sensor

Sun sensors provide the pitch and yaw position of the spacecraft

relative to the sun. It provides initial sun acquisition and backup or

emergency attitude knowledge during cruise mode. A CCD sensor has been

selected based on its redundancy, long life, reliability, high accuracy, low

mass and low power consumption (Flamenbaum, 234). Two units are used

for redundancy. One located on the scan platform boom and the other on

the HGA. They are aligned so at least one of them has the Sun in its field of

view at all times.

AACS Computer

The microprocessor is the heart of the AACS (Johnson, 4). All

measurements from the attitude sensors are processed by the computer,

which in turn gives the appropriate commands to the attitude control

hardware. The AACS computer contains the flight software and fault
protection algorithms to perform and maintain the AACS. The flight

software contains the algorithms to maintain Intrepid attitude and control

during all mission modes. The computer is also interfaced with the

communications subsystem allowing for reprogramming of the flight

software from Earth. The fault protection algorithms give the computer the

power to reconfigure the AACS in case of system problems. There are four

basic problem areas to be concerned; hardware failure, Single Event Upset

(SEU), flight software problems and ground errors (Johnson, 7). The fault

protection software continually monitors the health of subsystem, making

sure that it is functioning properly to maintain Intrepid attitude.

The microprocessor used is going to allocate 32k of the main

computer. It contains a bidirectional interface between it and the sensors

and actuators and is internally redundant. These are unique and

redundant sets of interfaces with each peripheral for optimum autonomy.

Reaction Control System

The reaction control system of the spacecraft produces the force

necessary to maintain the attitude of Intrepid. There are many different

ways of providing this necessary torque. The two different techniques

considered for use on Intrepid are reaction wheels (RW) and reaction

control thrusters (RCT).

When selecting which control system would be used, the basic overall

requirements were the driving factors once again. The RCT provided the

best maneuvering capability of the vehicle for emergency situations, delta-v

corrections and in handling any changing mission requirements. It also

provided the simplest system of operation.
The drawbacks of the RCT are the plume contamination created by

the hydrazine thrusters, the additional torques created, and the moderate

pointing capabilities in comparison with the reaction wheels. The moderate

pointing accuracy is justified by the use of the HPSP for instruments with

stringent pointing requirements, and the fact that the system is not used

during fly-bys. Plume contamination is also created by the RW system since

it also uses thrusters to unload the reaction wheels.

The RCS consists of twelve 0.2 N thrusters used on the Mariner Mark

H spacecraft (Bell, 796). The system is broken down into three sets of four

thrusters, with any set being able to perform the necessary attitude

maneuvers for autonomy. Appendix A.6.1 calculates the minimum size of

thrusters needed to perform a difficult maneuver.There is one set around

the spacecraft bus, the thrusters are 90 degrees apart, and the two other

sets are on pods around the truss of the propellant tank. The thrusters are

angled so each set can perform the necessary maneuver for any axis. A

layout of the thrusters around the bus is shown in figure 6.2.

Mission Modes

Intrepid is going to have to be able to perform in many different

mission modes. It is not going to have to perform at peak performance for

long time periods. There are going to be periods of maximum pointing

accuracy and stability, and there are going to be times of relatively easy

pointing requirements. The following are descriptions of all of the different

modes.
Separ_tiQn frQm 10unch v_hicl_

The AACS must stabilize the spacecraft after release from the

launch vehicle and make corrections for the proper launch trajectory. The

AACS should stabilize the vehicle's three inertial axes at a rate of 0.1

deg/sec (Dougherty, 19).

The spacecraft is going to spend most of the mission in cruise mode.

Cruise mode consists of keeping the high gain antenna (HGA) pointed at

Earth and correcting for trajectory errors. The HGA has two different

pointing modes. The first mode is for when the uplink/downlink is on.

This mode is going to operate for 8 hours at a time once every week and

maintain a pointing accuracy of 8.7 mrad (Bell, 799). Tl_e other mode of

cruise is when the uplinlddownlink is off. This mode is going to operate at

all other times except during trajectory correction maneuvers. It has to

maintain a pointing accuracy of 17.5 mrad within Earth (Bell, 798). The

cruise mode corrects for trajectory errors every two to three weeks. This is

commanded through the AACS microprocessor and the torque produced is

going to vary in magnitude from correction to correction.

During cruise, the attitude reference is going to be performed

primarily by the star tracker. The star trackers are also used to update the

IRU. The IRU's primary responsibilities are during thruster activation

and back-up attitude reference in case of a loss of any of the other sensors.

The Sun sensors are used for back-up sensing and for reacquiring stars in

case of a loss of star lock. The torque is produced by the 0.2N thrusters as

commanded by the microprocessor.
y (yaw)

r=.9m

.2 N thrusters

v

x (pitch)

z (roll)

(x-y plane)

Figure 6.2 Bus thruster layout
Planetary _n_Q_nter

The planetary encounter mode requires the most strict stability

during the entire mission due to the imaging constraints. The far

encounter phase pointing of the planetary fly-by consists of pointing the

imaging camera at the planet. This is going to be accomplished through

articulation of the HPSP. Thus the Intrepid-Earth communication link is

not broken during the far encounter phase since the camera is actuated on

the HPSP which is dynamically isolated from the bus. The near encounter

phase pointing consists of performing the UV-spectrometer scanning in

addition to maintaining the imaging. The planet is acquired using the

ASTROS II in a closed loop pointing scheme.

Right before the near encounter, the IRU is to be updated by the

celestial sensors and the RCT is shut off. The only attitude to be kept is on

the HPSP. The RCS is shut off so not to contaminate any of the scientific

equipment (Coupe', 30). The communication link is not kept during this

phase. Once the phase is over (the UV-spectrometer scanning is complete),

the celestial sensors reacquire the Sun and guide stars based on the attitude

determined by the IRU. Intrepid now returns to the far encounter phase to

complete the imaging of the planet.

During imaging, the AACS needs to create 2.43 (N m sec)/ sec for

spacecraft stability during imaging. This is the momentum required per

second of operation (see Appendix A.6.2) (Koepke, 37). A possible problem in

this area is maintaining this stability of the spacecraft during the near

encounter phase. The stability management is going to have to be studied

further to see the feasibility of this approach. A possible solution at an extra
mass and power cost would be to add reaction wheels to control the

spacecraft during this short encounter phase.

Delta-v burns

The AACS aligns the spacecraft on the correct trajectory line for the

delta-v burn, and must maintain thrust vector control (TVC) throughout

the burn. The TVC keeps the spacecraft pointing on the correct trajectory.

This is performed by pulsing the attitude thrusters as commanded by the

computer using input from all of the attitude sensors. A TVC pointing

error of 20 mrad is to be kept during the maneuver (Bell, 799). There is no

communication link during the thrust. Earth is reacquired at the

conclusion of the burn in the same manner as after a near planetary

encounter. The different modes are shown in figure 6.3.

Departure

.from Earth

launch _ Cruise -m,,-delta_ v

vehicle

Jupiter

Cruise "_" flyby

Pluto

_Cruise -m,-flyby -_-Cruise

Figure 6.3 Mission modes

Conclusion

The AACS developed for the Intrepid mission stresses reliability,

autonomy, low cost and simplicity. This has been accomplished while being

able to meet extreme pointing and stability requirements. Thus the system

is able to perform flawlessy while being able to also perform autonomously

at a low cost.
Appendix A.6.1

Turn spacecraft 90 degrees in 18 minutes, from Hubble Telescope

(Dougherty, 17).

Equation:

Oc = (2*F*L)*(V2)2/(Iv)

Solve for F, using: L = .65m

t = 1080 sec

Iv = 1808.8 kg m 2

with full tank (Koepke, Inert)

ec = _/2 rad

then F =.0054 N

Thus 0.2 N thrusters are going to more than enough to handle the

spacecraft.

Appendix A.6.2

The momentum per second required for maintaining stability during

planetary encounter imaging (Koepke, 37) (Kohlase, 35).

Ht=(F*ton) 2 * L 2 / (et * Iv) where F =.2N

ton=15 sec

L = .65 m

@t = 1.75"mrad

Iv = 1719.016 kg m 2

with 68 kg in tank (Koepke,Inert)

then Ht=1.264 N m sec/sec
Bell, Charles and David Lehman. "Mariner Mark H: Spacecraft

Control for the 1990's and Beyond." AAS 86-285, 1986.

Cheng,G. J. and G. Tracy. "The Mars Observer: Attitude and

Articulation Control Subsystem." AAS 88-001, 1988.

Coupe',G.M. and P.J. Witteveen, "The Attitude and Orbit Control System

for Giotto, ESA's Halley Encounter Mission." Automatic Control in

1982. pp. 27-34

Dougherty, H., C. Rodini and K. Tampertrine, "Space Telescope

Pointing Control." Automatic Control in Space 1982. pp. 16-27

Draper, Ronald F., JPL "The Mariner Mark H Program." AIAA-84-

0214, AIAA 22nd Aerospace Sciences Meeting Reno, Nevada,

December, 1984.

Flamebaum, S. and P. Anstett. "Multipurpose Sun Sensor Using CCD

Detector." IFAC Automatic Control in Space 1982.

Johnson, Stephen B. "Fault Protection Design for Unmanned

Interplanetary Spacecra_." AAS 87-001, 1987.

Lademan, Ernest. "Gyro Technology " AAS 87-020, 1987.

Olschansky, David. "Magellen AACS Closed Loop Testing." AAS 87-035,

1987.

Stephenson, R. "Galileo Attitude and Articulation Control System: A

Radiation Hard, High Precision, State of the Art Control System."

IFAC Automatic Control in Space 1985. pp. 83-90

Stokes, Loren. "Long-life Spacecraft Fiber Optic Rotation Sensor." AAS 87-

015, 1987.
Conclusion

Final considerations for the Intrepid spacecraft include an

implementation plan. Once the optimal design is finalized, a prototype

must be built and thoroughly tested. Further trade studies, sensitivity

analyses, and redesign may be required for increased optimization. Upon

successful completion of testing, the remaining three spacecraft will be

built. Since Intrepid is constructed mainly of off-the-shelf hardware, the

development and manufacturing timeframes are greatly reduced.

Of utmost importance when determining a design freeze date is the

incorporation of the latest computer and Plutoian system knowledge.

Significant gains can be realized by integrating the most recent computer

developments in the areas of self-autonomy, flexibility, and survivability.

Any discoveries concerning the Plutoian system should affect the mission

science objectives in order to retrieve the most meaningful and vital

information.
Appendix B- Mass and Power Requirements

Subsection

SI

SUBTOTAL

MMPC

Component

SUBTOTAL

CRD

PLS

MAG

USO

SSI

UVS

N/A

HGA

Bus structure

Outboard shear plate

Inboard shear plate

Science boom

RTG boom

Fasteners

Connector bracket

Shunt radiators

Tank support

Truss Adapter

Tank struts

RTG launch support

Misc. supports

MLI - bus

MLI - science

MLI - propulsion

Heaters - bus

Heaters - science

Heaters - propulsion

Louvers - bus

Louvers - science

Service valves (16)

Re6ulator valves (4)

Filter

Tubes, fittings,

sensors, etc...

100 N thrusters (4)

Mass [k_]

7.0

10.0

6.0

2.0

31.0

5.0

61.0

N/A

35.72

24.58

17.59

8.11

7.64

6.69

1.50

2.17

6.88

4.24

8.82

0.66

3.73

8.67

22.52

3.01

1.99

0.75

1.21

1.80

2.94

0.78

Power Rec 1. [W]

5.0

8.0

6.0

5.0

26.0

6.0

_0

N/A

N/A

tv

I!

Jt

gl

v!

_t

Iv

11

Iv

oO

vv

tv

6.0

7.0

7.0

t*

f!

6.80

2.67

0.17

4.44

170.00 20.0

3.68 N/A
Subsection

_AL

CCC

SUBTOTAL

AACS

SUIFIDTAL

ConanVenc

S/C TOTAL

Component

0.2 N thrusters (12)

Mass [k$]

7.20

Power RecI. [W]

_t

Propellant tank 16.10 "

Pressurant tank 16.15 "

He pressurant 2.24

N2H2 propellant

Power distribution

unit

2.4 kHz inverter

Power control unit

Shunt re_dator

Pyro switching unit

Modular RTG's

Pyrotechnics

X-band transponder

Solid state amplifier

Input isolators (2)

Output isolators(2)

Duplexers (2)

Receiver RF switch

Transmitter RF

switch

Low pass filter (2)

X-band attenuator

800.00 "

5.41 "

4.29 "

fB 5.65

3.99

4.01

34.06

5.00

921236

5.00

5.50

0.35

0.25

2.94

0.90

1.04

0.06

0.04

Command detector (2) 4.0

TLM Modulation (2) 4.0

Computers (2)

Solid State Mass

Memory

Power Supply

Misc. electronics

10.0

10.0

1.0

4.0

tt

vg

tt

N/A

Antenna subsystem

System assembly

5.0

5.0

59.08 90.0

FORS 10.0 10.0

ASTROS II 8.0 11.0

Sun sensors (2) 2.3

HPSP 11.81

33.11

1245.05

7.0

32.0

60.0

30.0

25_0
Appendix C - Trade Studies

RFP Requirements

Design Choices
CRS

PLS

SSI

A1-7000 T.M.

Batteries

Bipropellant

Cryog. prop

GN2 press.

He press.

Modular RTG

Monoprop.

Non mod RTG

One prop tank

Re_. Power

Solid prop.

Ti-6A14V tank

Titan IV

Titan IIID

Two prop tank

Unreg. power

Beryllium

GraphiteEpoxy

Passive Therm

Titanium

Toroidal Bus

Avail

1999

V/G

Withstand

Envir

V/G

Exist.

Hrdw

V/G

Fits

Miss.

Req.

V

Life

time

Low

cost

Recent

tech.

Reliable

V/G V/G

V/G V/G V/G G V/G G V/G

V/G V/G V/G G G V/G

X

V/G

X X X X X X X

Simple

X

X X X X X X X

X X X X X X

X X X X X X

X X X X X X X X

X

X

X

X

X

X

X

X

N/A

N/A

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

N/A

N/A

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

N/A

N/A

X

N/A

N/A

X

X

X X X X X X X X X

HGA X X X X X X X X

LGA X X X X X X X

X X X

X X X X

X X X X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

LowDataRate

Prog. in C

SA 3300

'SSA

_X-band only

'X-band transp.

ASTROS II

CCD Sun Sens

FORS

HPSP

Thrusters

X

X

X

X

X

X

X

X

X

X

X

X

X X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X
Appendix D - Acronyms

AACS

AU

BSTC

CCD

CCS

CDS

CM

CRAF

CRS

DSN

DSS

DTR

ES

FDS

FPA

HGA

IDC

JPL

LGA

MAG

MIPS

MLI °

MMII

MPS

PLS

PPS

PTC

RAM

RFI

RFP

RFS

RS

RTG

SA3300

S/C

SDS

SI

SNR

SS

SSA

SSI

SSP

USA

USO

Attitude and Articulation Control Subsystem

Astronomical Unit

Bus and Science Thermal Control

Charge-Coupled Device

Computer Command Subsystem

Command and Data Subsystem

Center of Mass

Comet Rondezvous Asteroid Fly-by

Cosmic-Ray detector System

Deep Space Network

Data Storage Subsystem

Digital Tape Recorder

Expert Systems

Flight Data Subsystem

Fault Protection Algorithm

High Gain Antenna

Image Data Compression

Jet Propulsion Laboratory

Low Gain Antenna

Magnetometer

Millions of Instructions Per Second

Multi-Layer Insulation

Mariner Mark H

Micrometeorite Protection System

Plasma detector System

Power and Propulsion Subsystem

Propulsion unit Thermal Control

Random Access Memory

Radio Frequency Interference

Request For Proposal

Radio Frequency Subsystem

Reed-Soloman

Radio-isotope Thermal Generator

Sandia Application 3300 Family

spacecraft

Structural Design Subsection

Scientific Instrumentation

Signal to Noise Ratio

Stuctural Subsystem

Solid State Amplifier

Solid State Imaging

Science Scan Platform

Upper Stage Assembly

Ultrastable Oscillator