N91-16046

Group 1

AAE Final Design P_port

April24th, 1990

- p / _.,) f_:!!>_

PFERD

Mission

Pluto Flyby Exploration/Research Design

https://ntrs.nasa.gov/search.jsp?R=19910006733 2019-05-09T14:34:54+00:00Z
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The PFERD mission will consist of a flyby spacecraft to the

planet Pluto and its satellite, Charon. The mission lifetime is

expected to be 18 years. The Titan IV with a Centaur upper stage

will be utilized to launch the craft into the transfer orbit. Each

subsystem of the craft was designed by a different individual and is

presented in a seperate section of the report. The group did tradeoff

studies to optimize all factors of design, including survivability,

performance, cost, and weight. Problems encountered in the design

were also presented.
f" 

, 

GROUP 1 • PFERD 

~fvtJ~~ Gary Lemke 

Group Leader 

Power and Propulsion 

~/H.¥ Husni Zaye~ 

AAC 

n Herring 

MMPC 

, 

Df~!iL Structures 

duM, );;;tfj;L __ 

Kevin Sutton 

SI 

, 

"V~L{Le S L~~v \" 

Mike Sharkey, ~ 

CCC 
TablQ Qf Contents

Abstract

Introduction

Scientific Instrumentation

Command, Communications, Control

Attitude and Articulation Control

Power and Propulsion

Structures and Thermal Control

Mission Management and Costing

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1

Introduction

The PFERD mission will be one of immense scientific interest.

Since its discovery, not much knowledge has been gained about this

far away planet. It is the purpose of this mission to change this.

Our mission has been dubbed PFERD, which stands for Pluto

Flyby Exploration/Research Design. It will consist of a Pluto flyby

spacecraft and all of the components needed to send it to Pluto.

Our proposal has been divided into six main subsystems. They

are, in order of appearance in this paper, Scientific Instrumentation;

Command, Communications, and Control; Attitude and Articulation

Control; Power and Propulsion; Structures and Thermal Control; and

finally, Mission Management and Costing.
SCIENTIFIC INSTRUMENTATION

by KEVIN L. SU'B'ON

2

The ultimate goal of this mission is the return of new

scientific information about Pluto and its satellite Charon. The

discovery of Pluto occurred in 1930 while Charon was not

discovered until 1978. During the six decades following the

discovery of Pluto, determining the characteristics of the planet has

been a difficult endeavor. Although Pluto was at perihelion in 1989

(29.6 AU), Pluto's mean distance from the sun is 39.5 AU. Pluto, s

orbital period is 248 years. The physical parameters of the PlutoCharon system have been derived from mutual event observations of

the system from 1985 through 1988. Table S1-1 lists the values of

the most extensive analysis of the mutual events to date.

TABLE Sl-l: Pluto-Charon.Physical Parameters (Binzel, 1989)

Semimajor Axis 1 9640 +/- 320 km

Eccentricity 0.0001 +/- 0.001

Period 6.387245 +/- 0.000012 days

Pluto's Radius 1150 +/- 7.0 km

Charon's Radius 593 +/- 10.0 km

Mean Density 2.030 +/- 0.035 gm/cm^3
3

The individual densities of Pluto and Charon cannot be determined

because the mutual event observations cannot predict individual

densities. There are many other uncertainties about Pluto and

Charon, including the following questions:

t

o

Does methane frost cover the surface of Pluto ?

What composition and structure does the atmosphere

have ? What is the haze layer composed of ?

What is the composition and structure of the bodies ?

Are there color variations over the surface of Pluto ?

What is the origin of Charon ?

Is an atmosphere refreezing to the surface of Pluto as it

moves away from the sun and at what rate ?

Are there any other satellites or rings ?

What is covering Charon's surface ? Water frost ?

What is the nature of the magnetic field and the

interaction with the solar wind ?

What is the population of the proposed Kuiper Comet Belt

(30-50 AU from the sun)

SUBSYSTEM REQUIREMENTS

The Request For Proposal lists the general requirements of the

overall spacecraft design. Some additional requirements for the

scientific instrumentation subsystem include 1) describing and

justifying the science objectives, 2) selecting and optimizing the

instruments, and 3) determining the location, mass, power

requirements, and data rate of the selected instruments. These

requirements must be met while also stressing reliability,

simplicity, and low cost. Performance must be optimized while
minimizing the mass of the subsystem. Materials or techniques

expected to be available after 1999 cannot be used.

SCIENCE OBJECTIVES

This mission to Pluto should answer all of the questions about

the planet, in addition, many unprecedented discoveries should be

made. The science objectives have been determined so that all of

the true values for the many uncertainties will be revealed. The

science objectives of the mission are:

1)

2)

3)

4)

S)

6)

7)

8)

9)

Determine the composition and structure of Pluto's atmosphere

Determine the mass, composition, and structure of Pluto and

Charon

Determine the dynamics of the Pluto-Charon system

Determine the color variation over the surface of the planet

Determine the nature of the magnetic field

Determine the origin of Charon

Study the impacts and impact rates to estimate the population

and mass of the proposed Kuiper Comet Belt

Determine the interaction with the solar wind

Search for any satellites or rings
INSTRUMENT SELECTION

The instruments have been selected to accomplish the science

objectives of the mission. The first step of the selection process

was to examine existing or planned spacecraft to determine what

off-the-shelf instruments were available to help minimize costs.

The space vehicles researched include Voyagers 1 and 2, Galileo,

Magellan, Pioneer 10 & 11, Giotto, Mars Observer, microspacecraft,

and the Mariner Mark II program (CRAF and CASSINI). To meet the

science objectives, the following instruments are desired:

5

Solid state imaging system (SSI) take pictures to help investigate

the surfaces and atmospheres of the two bodies, the

magnetospheric interactions, the system dynamics, and

conduct other visual searches.

Photopolarimeter (PPO) - determines the distribution and

character of atmospheric particles (determines the nature

of the haze layer).

Ultraviolet Spectrometer (UVS) - measures gases in the atmosphere

to determine its composition and structure.

Infrared Spectrometer (IRS) - determines the composition and

structure of the surface of the planet and satellite.

Magnetometer (MAG) - monitors the magnetic field for strength

and changes.
Plasma Analyzer (PLA)- determines the interaction with the solar

wind.

Radio Science (RSC) - determines the dynamics of the system,

using the high gain antenna and the communications

equipment.

Gamma Ray Spectrometer (GRS) - measures the elemental

composition of the surface of the planet.

Laser Altimeter (LAT) - determines the global topography of the

planet.

6

Once it became clear that this mission would involve the flyby

of a spacecraft instead of an orbiting spacecraft, the laser

altimeter and the gamma ray spectrometer were eliminated because

they were designed for an orbiting spacecraft (Komro, 1989).

The choice of a specific imaging system involves many

decisions. The Voyager imaging system has been proven to be

reliable over long periods of time in space, although it uses outdated

technology. The Galileo imaging system uses charged-coupled

devices allowing for advanced solid state imaging. The imaging

system designed for a microspacecraft is very light weight, but it

does not give good resolution. Table SI-2 gives a comparison of the

three imaging systems.
?

TABLE Si-2: Imaging System Comparison

(Flight 1987, Galileo 1985, Jones 1989)

Voyager Galileo "micro"

Mass (kg) 30 28 0.8

Power (W) 29 17 3.8

Resolution 0.07 m/pixel 0.07 m/pixel 7.0 m/pixel

@ 1000 km @ 1000kin @ 100km

At first, the microspacecraft camera seems to be the best. It

is very light weight and consumes much less power than the other

two systems, but its resolution is much worse than the other two

systems and it has not been proven in space. Due to these negative

factors of the microspacecraft camera, it was not given any further

consideration for use. Between the Galileo and Voyager imaging

systems, the Galileo system represents the best choice, because it

uses the latest imaging technology, has the least mass, and

consumes the least amount of power, so it will be included on the

Pluto probe.

To achieve reliability and low costs all of the instruments to

be included on the probe are existing instruments from other

spacecraft systems. Table SI-3 gives the mass, power

requirements, data rate, and spacecraft of origin for each of the

instruments to be included on the spacecraft.
8

TABLE Sl-3: Instrument Characteristics

(Flight 1987, Galileo 1985, Report 1985)

INSTRUMENT MASS DATA RATE POWER ORIGIN

(kg) (kbps) (W)

SSI

PPO

UVS

IRS

MAG

PLA

RSC

28 115.2 17 GALILEO

5.0 10.0 10 VOYAGER

4.0 0.1-2.0 5.3 GALILEO

18 0.5-10.0 12 CRAF

3 0.01-0.4 4 CRAF

5 0.01-115.2 4 VOYAGER

....... GALILEO

TOTALS 63.0 144.6-271.6 52.3
9

INSTRUMENT LOCATION

The instruments have specific requirements that specify

where they can be located. Some of the instruments like the imaging

system, need and unobstructed field of view. The high gain antenna

is the main source of obstruction. To give a good field of view the

instruments will be mounted on a high precision scan platform. This

scan platform will be located on a boom and have two degrees of

freedom so that it gives the instruments located on it an almost

unobstructed field of view in any direction. The scan platform

requires a pointing accuracy of 0.0034 rad and a slew rate of .00576

rad/sec to accommodate the instruments. The magnetometer needs

to be located far away from the electronics bus because if the

electronics bus generates a magnetic field, it will interfere with

the magnetometer's sensors. To minimize this problem, the

magnetometers will be located on an extendable boom of their own.

The boom, when extended, is eleven meters long. One magnetometer

sensor is located at the end of the boom while another is located is

located approximately five meters from the end of the boom. The

magnetometer electronics are located in the electronics bus in order

to isolate the magnetometer sensors. The location of each of the

instruments is given in Table SI-4. Figure S1-1 is a scale drawing of

the high precision scan platform and the instruments that are

located on it and Figure SI-2 is a scale drawing of the extendable

boom and the magnetometer sensors.
10

TABLE Sl-4: Instrument Location

High Precision Scan Platform Imaging System

Ultra Violet Spectrometer

Infrared Spectrometer

Photopolarimeter

Plasma Analyzer

Extendable Boom Magnetometer Sensors
FIGURE S1-1 ' High Precision Scan Platform

_PLA

AACS Se,15ors

]

O. 3"75

UVS

PPQ

IRS

SSI

FIGURE SI-2 • Magnetometer Boom
CONCLUSION

The science instruments have been selected to maximize the

scientific return for a flyby mission to Pluto, while also minimizing

weight and cost. Off-the-shelf instruments have been incorporated

into a scientific package that is simple, yet reliable. The

instruments will meet or exceed all of the objectives of the

mission.
L3

REFERENCES

Binzel, Richard P., "Pluto-Charon Mutual Events', Geophysical

Research Letters, Vol. 16, No. 11, November 1989,

p.1205-1208.

"Flight Science Office/Science and Mission Systems Handbook',

JPL D-498, October 1, 1987.

*Galileo: Exploration of Jupiter's System', NASA SP-479, June

1985

Jones, Ross M., "Think Small In Large Numbers', Aerospace

America, October 1989, p. 14-17.

Komro, F., "The Mars Observer Instrument Complement', AIAA

PAPER 89-0258.

"Report of the Comet Rendezvous Science Working Group', NASA

Technical Memorandum 87564, April 1, 1985.
Command, Control and Communication

In order to determine the essential requirements for the

Command, Control and Communication Subsystem, the Request For

Proposal document must be examined. These requirements were

found to be as follows: optimize performance, minimize weight and

cost, use off-the-shelf and reliable hardware, materials and

techniques must be developed before 1999, the design lifetime has

to be sufficient for the mission with a safety margin, the spacecraft

must communicate a distance of 38 A.U.'s and the subsystem can not

conflict with the other subsystems. What separates this mission

from all previous missions is the great distance the spacecraft must

travel. The challenge presented to the C.C.C. Subsystem is the

ability to communicate with Earth at this distance. Therefore this

report focuses mainly on communication. The other aspects of C.C.C.

are covered, but to a lessor extent due to the fact that they will be

more standard and similar in design to previous missions.

The main part of communication is the choice of antenna to be

used. In comparing the parabolic vs. the isotropic antenna, it can be

shown that the parabolic antenna produces almost thirty thousand

times the power received back at Earth of that produced by the

isotropic antenna ((Yuen, p. 6) eqn.#1). This is because the isotropic

antenna radiates in all directions while the parabolic antenna
concentrates its waves in a cone configuration. Therefore a

parabolic antenna is selected over an isotropic antenna.

Now that the parabolic antenna has been selected, "to achieve

best possible performance, we must design the telecommunications

system which gives the highest signal power, lowest amount of

noise, and most efficient use of signal-to-noise ratio, within

constraints such as spacecraft weight, size and cost (Yuen, p.3)." We

want to optimize the power received back at Earth. Looking at the

equation for the power received (Yuen, p.6), there are several ways

to increase the power received. These are: increase the

transmitting power, increase the diameter of the receiver, increase

the diameter of the transmitter, decrease the wavelength used and

decrease the transmitting distance. Decreasing the transmitting

distance might entail putting some sort of transmitter half-way

between Pluto and Earth. But this would be another mission in itself

and is not considered an option. Next we can look at increasing the

power transmitted, but this will be a set amount depending on how

much power is available from the Power Subsystem. "Spacecrafttransmitted power is typically only 20 watts (Yuen, p.4)." There are

a couple of ways to increase the diameter of the receiver. The most

obvious being to make a larger and larger receiver. But this is too

costly. A second method involves arraying already built receivers

electronically to increase the effective area of reception. "This

network is being upgraded to nine antennas: six 34-meter antennas

and three 64-meter ones...the DSN 64-rneter antennas will be

enlarged to 70 meters (Posner, Horttor and Grant, p.62)." By arraying

these antenna receivers, the power received will be increased by
more than 5 times that of just one 64-meter receiver (eqn. #2).

Arraying the receivers into a network is a good option. By doubling

the diameter of the transmitter on the spacecraft, this will increase

the power received by 4 times (eqn. #3). But the weight of the

antenna will be doubled, and that does not sit well with the Mission

Management Subsystem who is trying to minimize the weight. Also

"the largest planetary spacecraft antenna yet is 4.8 meters (Posner

and Stevens, p.20)," meaning a bigger antenna would not meet the

R.F.P. requirements of off-the-shelf reliable material developed

before 1999. Therefore increasing the transmitter diameter antenna

greater than 4.8 meters is a bad option. Lastly, we can decrease the

wavelength used in transmitting. To do this we must increase the

frequency used. There are assigned frequencies used for space

communications so that outside interference is minimized. The Xband (8.4 GHz) is now the standard down-link frequency used (Posner

and Stevens, p.8). But by 1995, Uthe down-link could well be at 32

GHz (Posner, Horttor and Grant, p.62), u By using this Ka-band, it can

increase the power received by almost 15 times of that using the Xband (eqn. #4). Therefore using the Ka-band (32 GHz) frequency is an

excellent way to increase the power received.

Another method for transmitting data to Earth is through laser

technology. In comparison with the 20 watts needed for the

parabolic antenna, the laser only needs .5-2 watts of power and is

only 10 cm. in diameter (Lesh and Rayman, p.81). This gives the

laser a great weight reduction advantage. Also because the

wavelength of a laser is only .5 micro-meters, this increases the

efficiency of the signal at Earth one million fold compared to the
17

parabolic antenna (Lesh, p.106). And the laser can provide all sorts

of new "light sciences (Lesh and Rayman, p.84)." But the laser also

has disadvantages at this time. The laser must be pointed with

extreme accuracy. "With the long propagation time, you only have

one shot at beam acquisition (Lesh, p.106)". Also little deep space

testing with lasers has been done, which means its reliability is

unknown. We do not know if the laser technology will be complete

for deep space use by 1999. Therefore, laser technology is in direct

conflict with the R.F.P. and can not be used.

After analyzing all the options for communication with Earth, we

selected the parabolic antenna to be the best. The diameter will be

4.8 meters (the largest spacecraft antenna available) to optimize

the power received. The network of the nine receiving antennas

arrayed together will be used also to optimize the power received.

The wavelength will be decreased by using the Ka-band (32 GHz)

frequency to increase the power received. The 4.8 meter diameter

parabolic antenna, along with the arrayed receivers and Ka-band

frequency, will give the spacecraft the best possible power received

at Earth while staying within the constraints of the R.F.P.

The spacecraft C.C.C. Subsystem must provide the Scientific

Instrument Subsystem with a maximum data rate estimate so that

the S.I. Subsystem can know what amount of data he will be able to

send to Earth. The data rate is mainly dependent on the signal-tonoise ratio and the power received. Assuming a signal-to-noise

ratio of 20, the power received can be calculated using the parabolic

antenna and options chosen earlier (eqn. #5). The power received

equals 1.593"E-16 watts. Therefore a data rate estimate for the S.I.
18

Subsystem equals 316891 bits/sec (eqn. #6). If this data rate is not

large enough for the S.I. Subsystem, then storage considerations for

the data are necessary.

optical storage disk.

being transmitted. "With

compressed imaging output

received with a very low

compression, a single error

(Posner, Horttor and Grant, p.63)."

bit error probability;

can destroy a large

This can best be achieved through use of an

Another possibility is to compress the data

Galileo, this can raise the partly

rate...400 times...Such data must be

because of the

amount of data

In order to communicate with Earth, the spacecraft antenna must

be directed toward Earth. "The sun is still the primary attitude

reference (J.P.L., p.19)," which is used to point the antenna toward

Earth. The antenna will be mounted on the front of the spacecraft to

avoid the delta-v burns used to go to Pluto. Therefore, the Structure

Subsystem must provide a shield for the antenna to combat

environmental and atmospheric hazards. After the initial delta-v

burn, the Attitude and Articulation Subsystem must rotate the

spacecraft 180 degrees so that the antenna is facing the Earth. If

another delta-v burn is necessary, the spacecraft must first be

rotated 180 degrees back into its proper position. Then after the

burn is complete the spacecraft can be rotated 180 degrees once

again to face the Earth. This process will need to be repeated as

many times as the number of delta-v's necessary for the mission.

For the Command and Control part of C.C.C. Subsystem, we must

look at the use of computers on-board the spacecraft. One problem

with the distance that must be traveled for this mission is that it

takes over five hours for a signal sent from Earth to reach Pluto
L9

(eqn. #7). In a time of crisis aboard the spacecraft, five hours may

be too long of a time to wait for a command. Therefore, it is

necessary that Artificial Intelligence be available to the spacecraft

so that it can analyze a situation and make a decision on its own to

correct the problem. As far as the type of computer system to be

used, one similar to that used for Galileo or Voyager would be a good

choice because of their proven reliability. The only problem with

these computer systems is that they are ancient. They are very slow

and their memory capabilities are limited. Therefore, we selected

the advanced High Performance Micro Computer. This computer

contains a 2 million Byte memory, uses 20 watts of power, and only

weighs .1 kg ( Jones, p.11). Also, the Command and Control

Subsystem "can survive any single internal fault, because each of its

functional units has a duplicate elsewhere in the subsystem (J.P.L.,

p.21)." Therefore with the Command and Control Subsystem

completed, this concludes the design for the Command, Control and

Communication Subsystem.
20

#1

A.Doendix 1: Eauations

Parabolic vs. Isotropic Antenna

Parabolic Power Received (Pr)

Pr = Pt*Lt*Gt*Ls*Lr*Gr

Gt = .55"SQR(3.14159*Dt/Wavelength)

Ls = SQR(Wavelength/12.56*r)

Gr = .55*SQR(3.14159*Dr/Wavelength)

Isotropic Power Received (Pri)

Pri = .5*Ar*Pt/12.56*SQR(r)

Assume: Pt = 20 watts, Lt = Lr = .5, Dr = 64 meters

r = 38.5 A.U., Dt = 4.8 meters,

Wavelength = 3*E8(m/s)/8.4(GHz)

W = .0357 meters

Pr (para) = 2.082"E-18 W Pri(iso) = 7.716"E-23 W

Pr(para) = 26983*Pri(iso)

#2 Increase the diameter of Receiver

Dr = SqrRoot(6*SQR(34) +3*SQR(70)) = 147 meters

Pr = (SQR(147)/SQR(64))*Pr(original)

Pr = 5.28"Pro

#3 Increase the diameter of Transmitter

Pr = (SQR(2*4.8)/SQR(4.8))*Pro = 4*Pro
23.

#4 Decrease the Wavelength

Wavelength -- 3*E8/Frequency

Freq. - 32GHz

Pr- (SQR(32)/SQR(8.4))*Pro = 14.5 Pro

#5 Power Received for Parabolic Antenna with Options

Pr- Pt*Lt*Gt*Ls*Lr*Gr

Pt - 20 W

Lt - Lr = .5

Gt - .55"SQR(3.14159*4.8/(3*E8/32GHz))

Ls = SQR((3*E8/32GHz)/12.56*5.76*E12)

Gr = .55"SQR(3.14159*147/(3*E8/32GHz))

Pr - 1.593"E-16 W

#6 Data Rate (B)

B = w*log(SNR + 1)/Iog(2)

w = Pr/k*T*SNR

k =, 1.38*E-23(J/K)

T = 8K

SNR = 20

B = 316891 (bits/sec)

#7 Time of Transmission to Pluto

Time = distance/Velocity

dist. = 5.76"E12 meters

vel. = 3"E8 meters/sec

Time = 5.33 Hours
Reference Paae 22

Lesh, James R. ;

1988

Rayman, Marc D.

"Deep-space missions look to laser communications"

(Oct.): p.81-2.

Lesh, James R.

1988 "Laser-based systems will enhance deep-space

optical communications" _ (Feb.)" p.106.

Posner, Edward C. ;Stevens, Robertson

1984 "Deep space communication--past, present, and

future" IEEE Communications Maaazine (May)" p.8-21.

Posner, Edward C. ; Horttor, Richard L. ; Grant, Terry L.

1986

Yuen, Joseph H.

1983

Jet Propulsion Lab

1989

Jones, Ross M.

1990

"Communicating from the edge of the solar system"

IEEE S0ectrum (June): p.60-5.

Deeo Soace

p.1-6.

Telecommunications Svstems Enaineerino

"Flight science office • science and mission systems

handbook" Voyager Neotune / Interstellar Mission

(Oct. 1): p.19-21.

"Microspacecraft technology for space science

missions" Ca. Inst. of Technology/JPL (Jan. 8): p.11.
23

Attitude and Articulation Control Subsystem.

We can now examine the Attitude and Articulation Control

System (AACS) for our spacecraft. The Request for Proposal (RFP)

requires that the AACS design should 1) optimize weight, cost and

performance, 2) be reliable and easy to operate, 3) use off-the-shelf

hardware when possible, 4) be able to have a lifetime sufficient to

carry out the mission plus a safety margin and 5) be able to perform

several possible missions. The mission itself required that the

system should guarantee communications with Earth, maintain the

spacecraft's trajectory and be highly autonomous due to the mission

length and the distance the spacecraft must travel away from Earth.

These requirements served as a guideline in the design of our

spacecraft's AACS. The primary AACS hardware selected consists of

a star tracker, a gyroscope, a sun sensor, a computer and an

assembly of thrusters for the attitude and trajectory correction

maneuvers as well as the corresponding electronics and actuators to

complete the system.

High Precision Scan Platform Sensors.

The star tracker selected for our mission is the Advanced Star

and Target Tracking Optical Sensor (ASTROS II), and the gyroscope

selected is the Fiber Optic Rotation Sensor (FORS). Both of these

sensors were selected from the Mariner Mark I1: Comet Rendezvous

and Asteroid Flyby (CRAF) mission (Bell and Lehman). The ASTROS II

was selected because it enables closed loop target tracking which

allows for autonomous science data gathering (Bell and Lehman). It
24

is also relatively lightweight, with a mass of 11 kg, and has a

relatively low power requirement of 15 Watts (Bell and Lehman). The

ASTROS II also provides very accurate star tracking to 20 arcsec for

up to three stars simultaneously and allows for the autonomous

calibration of the gyroscopes based on the star tracker data (Bell

and Lehman). FORS provides low mass, solid state inertial angular

rate and position sensing and is designed to meet or exceed NASA

DRIRU II performance specifications (Bell and Lehman). These two

instruments were also selected because they exceed the the

requirements imposed by the Science Instrumentation subsystem.

This subsystem required that the camera have a pointing accuracy of

.0034 radians and that the scan platform have a slew rate of .00576

radians per second. The pointing requirement is met by ASTROS II

which provides a target dependent accuracy from 1 to 10 arcsec and

the slew rate requirement is met by FORS which provides a slew

rate range from .00523 radians per second to .06981 radians per

second (Bell and Lehman). The High Precision Scan Platform (HPSP)

was selected for the placement of these sensors for several reasons.

It provides an adequate separation distance from the contamination

of the attitude thruster exhaust and from the radiation generated by

the Radioactive Isotope Thermoelectric Generators (RTG). The HPSP

was also selected because it minimizes any translational and

rotational errors between the sensors and the science instruments

(Bell and Lehman).

Bus Sensors and Hardware.
25

The inertial attitude as determined by ASTROS II and FORS is

transferred to the basebody of the spacecraft to provide bus inertial

rate, position knowledge for High Gain Antenna (HGA) pointing and

thrust vector control (Bell and Lehman). This attitude determination

is backed up by the a Fine Sun Sensor Assembly (FSSA) to provide

redundancy. The FSSA was selected because it was used on the GRO

satellite and is thus flight proven and reliable and also because it

provides lightweight, low power redundancy with a mass of 1.75 kg

and a power input of approximately 3.5 Watts (Wertz). It was also

selected because it meets the pointing accuracy requirement

imposed by the Communications subsystem of approximately .15

degrees to .50 degrees by providing an accuracy of .022 degrees

(Jerkorsky, Keranen, Koehler, Tung and Ward).

An onboard computer (OBC) was needed to handle the autonomy

required by the mission and the storage of science and

communications data as well as the implementation of the attitude

correction maneuvers (ACM) and trajectory correction maneuvers

(TCM) determined by the sensors. To accomplish this task a high

performance micro computer, developed by the Lawrence Livermore

National Laboratory, will be placed in the bus of the spacecraft.

This computer was selected because it is extremely lightweight, has

relatively low power requirements and also because its storage

capability of 2 million bytes is over 50 times more powerful than

the computers used in the Galileo and Voyager missions (Koepke). A

total of 2 computers will be used to provide full redundancy even

though only one computer will operate at a given time.
26

AACS Pro oulsion System.

To perform all ACM's, TCM's and gravity assist maneuvers

(GAM) required, thrusters were chosen over reaction wheels.

Thrusters were selected because they 1) provide easier and quicker

rotation of the spacecraft due to its large dry weight of 500 kg, 2)

have been used on many other missions and are therefore reliable

and 3) provide enough accuracy in combination with the attitude

sensors for all science instrumentation. I studied several possible

AACS propulsion systems that would handle the requirements

imposed by the mission subsystems. A description of each is given

below along with with the reasons for disqualification or

acceptance.

I). 12, 10 Newton thrusters shielded and mounted in sets of 6 on

booms protruding from opposite sides of the bus for all ACM's and

TCM's, in combination with a 400 Newton engine used for all GAM's

(Yates, Johnson, Colin, Fanale, Frank and Hunten). This system is

identical to the system used on the Galileo spacecraft and is

therefore reliable, but the problem is that the system is fueled by a

bipropellant which will not last the duration of our mission of 18-

22 years.

II). 12, 10 Newton hydrazine fueled thrusters mounted and positioned

as described above in combination with a bipropellant fueled extra

complete stage used for all GAM's. This option was disqualified

because of its weight and high cost.

III). 24, 10 Newton thrusters in combination with a 400 Newton

engine. This system will be divided up into 2 sets. The first set will

contain 12 thrusters mounted and positioned as described above and
2"7

fueled by hydrazine and the second set will consist of the other 12

thrusters and the 400 Newton engine which will be fueled by a

bipropellant. The first set will be used for all ACM's and TCM's that

are required after all GAM's are completed. In the second set the

thrusters will be mounted on the 400 Newton engine. These along

with the engine will perform all ACM's, TCM's and GAM's needed from

the time of launch until the completion of the last GAM. Upon

completion of the last GAM the engine and its thrusters will be

jettisoned. This AACS propulsion system was chosen because it uses

the same thrusters and engine that were used in the Galileo mission

and because it uses bipropellant in the second set which has a better

specific impulse than hydrazine. The total delta V needed to be

generated by this system is approximately 3.3 km/s to 3.5 km/s and

has been estimated from the requirements imposed by the Mission

Management Planning and Costing subsystem (MMPC). The breakdown

of the delta V is estimated as follows. A delta V of 1.7 km/s is

needed for all GAM's to insure that the spacecraft will make it to

Pluto. This was determined by the MMPC subsystem. The delta V

required for all ACM's and TCM's is approximately 1.6 km/s and was

calculated assuming that the spacecraft needed a delta V of .12

km/s every 1.5 years for 20 years.
28

Charles Bell and David Lehman. "Mariner Mark I1" Spacecraft Control

for the 1990's and Beyond". Advances in the Astronautical Sciences.

1987. American Astronautical Society.

Jerkorsky, Keranen, Koehler, Tung and Ward. "GRO Attitude Control

and Determination". Advances in the Astronautical Sciences:

Guidance and Control 1986. 1986. American Astronautical Society.

JPL. Vovaa_er Neotune/Interstellar Mission. Ftiaht Sciences Office.

Science and Mission Systems Handbook. 1987. JPL.

Koepke. "Microspacecraft class notes". Spring 1990. Kinko's

Wertz. $0acecraft Attitude Determination and Control.

Reidel Publishing Company.

1987. D.

Yates, Johnson, Colen, Fanale, Frank and Hunten. Galileo: Exoloration

of Juoiter's System. 1985. Library of Congress.
29

POWER AND PROPULSION SUBSYSTEM

POWER SYSTEM

The main requirement for the power source is that it must

deliver enough reliable and uninterrupted power throughout the

lifetime of the mission with a sufficient safety margin. The other

requirements for the power subsystem are as follows. It must be

designed for reliability, simplicity and low cost. It must be easy to

operate and use off the shelf equipment where possible. All

technology must be available on or before 1999. The power delivery

system must protect the circuits, protect the load, and be able to

control and distribute the power.

When selecting from the possible power sources, the most

common space power source, solar, was eliminated from

consideration because it would not be able to produce enough power

at the distances that our mission would cover (over 40 AU).

Batteries and fuel cells would not have a useful lifetime sufficient

for our mission, so they were also eliminated from consideration.

The two sources of power that could supply power at 40 AU and

beyond for the duration of our mission are a space nuclear reactor

and a Radioactive Isotope Thermoelectric Generator (RTG). RTG's

were selected to provide the power for our mission because the have

been proven to be reliable, safe, and easy to operate on several deep

space missions. A space nuclear reactor was eliminated from

consideration because there are no current space qualified reactors,

and the only one that is currently being designed is designed to

produce 100 mW, which is approximately 300 times larger than our
3O

required power of 305.5 Watts. The RTG's are also much safer from

an environmental aspect, and they will require much less shielding

to protect the scientific instrumentation. This analysis is

summarized in table PP-1.

POSSIBLE POWER SOURCES

ADVANTAGES

sufficient life X

operate at 40AU X

fully developed X

technology

flown in space X

proven reliable X

on long duration

spacecraft

RTG NUCLEAR SOLAR

X X

X

X X

X

X

BAI-rERY FUEL CELLS

X X

X X

X

X

(if recharged)

X

Table PP-1. TRADE FOR POWER SOURCE
The type of RTG that we will utilize for our mission will be

similar to the design proposed in a study by Fairchild (Schock). This

design has many advantages over previous designs. This study was

done to optimize the design of current RTG's by incorporating the

latest developed power source, the newest materials, and utilizing a

modular design so that it would be able to be used for many

missions. The power system for our mission will be capable of

305.5 watts. The breakdown of the power requirements for the

various subsystems is shown in table PP-2. The RTG thermal power

source consists of 13 modular slices of the General Purpose Heat

Source (GPHS). Each thermal power source slice delivers 250 Watts

of thermal power, which will be converted into 23.5 Watts of

electric power. The Modular Isotopic Thermoelectric Generator

(MITG) design was selected over the two previously used RTG's these

possible RTG designs are compared in Table PP-3 The Multi Hundred

Watt system is the one that was flown on the voyager missions, so

it has been proven in flight, but it does not take advantage of any of

the recent improvements in RTG designs. The GPHS/RTG takes

advantage of the new modular General Purpose Heat Source, but it

does not utilize a fully modular design. It also does not make use of

the new thermoelectric materials (SiGe+GaP instead of just SiGe).

The only disadvantage to the MITG design is that it has not flown or

even been produced, but developing a new RTG based on this design

will more than double the power to weight ration of current RTG's

(Schock, p342). This RTG is Shown in Figure PP-I.
32

POWER REQUIREMENT BREAKDOWN

Scientific Instrumentation

Attitude and Articulation

Mission Planning

Command, Control, and Communication

Propulsion

Structures

15 % for lifetime losses

Total power required

Total power delivered (nearest modular power)

Table PP-2. Power Requirements

75 Watts

52 Watts

0 Watts

40 Watts

20 Watts

70 Watts

40 Watts

297 Watts

305.5 Watts
33

TYPES OF RTG'S

ADVANTAGES

apx. specific power

able to provide

required power

uses modular

heat source

fully modular

design

uses latest

thermoelectric

materials

flown in space

Table PP-3.

MITG

4.7 W/Ib

X

X X

X

X

X

TRADE FOR RTG TYPE

GPHS/RTG

2.3 W/Ib

X

MHW

1.8 W/Ib

X

X
34

31.5 in

Figure PP-1
35

Power Supply sizing

The size of the RTG power supply is based on the power

estimates listed in table PP-2. The 15% additional is to account for

the degradation of the thermal power source over the lifetime of the

mission. The weight breakdown for the total power supply system is

given in table PP-4. These weights were calculated by scaling the

weights of all of the components that will either be larger or that

there will be more of in a larger RTG design (such as the number of

modular heat sources), and then adding in the weights that would be

constant for any size RTG (such as the end plates and their

associated mounting hardware).
36

WEIGHTS

Housing

outer shell (1)

fins (4)

emissive coating

aux cooling manifold (1)

nuts (4)

6.5

2.5

0.27

0.325

0.07

Converter

TIE module (128)

T/E C-seals (128)

nuts (4)

foil ins. (1)

foil ins ends (2)

power converter (1)

gas management assembly (1)

electrical straps

PRD (1)

C-seal- ends (2)

other insulation

end caps (2)

screws - end caps (32)

pads and bushings

Heat Source Support System

load spreaders

PG buttons

Bushings

Belleville Springs

Pistons

Compression plates

Preload Screws

(8)

(8)

(8)

(16)

(8)

(40)

(8)

Heat Source (3250 Watts Thermal)

Heat source module (13)

Total Weight

Ibs

kg

4.12

0.0325

0.038

2.275

0.20

0.35

0.29

0.40

0.90

0.06

0.10

1.47

0.17

1.00

0.35

0 04

0 09

014

032

006

031

41.62

64.0

29.1

Table 4. Power Subsystem Weight Breakdown
Possible Use of Batteries

Batteries were considered for use in conjunction with the

RTG's for power. They were looked at to be utilized when there was

a peek demand for power. When analyzing the mission, this would be

when transmitting from the Vicinity of Pluto, (or other planetary

encounters). They could be charged on the way there and then

utilized when transmitting. This did not turn out to be such a good

idea. The batteries do not have that great of a weight advantage

over the RTG source, and it would interrupt the mission when the

batteries had to be recharged. Therefor, the added complexity for

the power subsystem and the interference with the mission

objectives ruled out the use of this type of hybrid power system.

37

Power conditioning and regulation

Because there are no batteries in the power system and

because the RTG's supply a fairly constant voltage, power regulation

does not seem to be a problem with RTG sources. The breakdown of

the thermal source will show up as a loss of current, and the voltage

will be basically constant. The wiring of the RTG will be redundant

to increase the reliability and to ensure that no catastrophic loss of

power will occur.
38

PROPULSION

The propulsion system for our spacecraft consists of two main

subsections. The first will consist of the thrusters and fuel for the

gravity assisted delta V at Jupiter and the thrusters and fuel for the

AAC maneuvers for the portion of the mission from Earth to Jupiter.

The second will consist of the thrusters and fuel for the AAC

maneuvers from Jupiter to Pluto and beyond.

FUEL SELECTION

Trade studies have been performed to select the optimum fuel

for the specific requirements of each stage. The Fuels considered

for the various stages are: monopropellant hydrazine, bipropellant

N204/MMH, bipropellant liquid oxygen and liquid hydrogen, and solid

propellant. The advantages of hydrazine are that it is a

monopropellant so it is simple and very reliable, and would therefor

costs less than other systems, but it has very low performance.

Bipropellant N204/MMH has much better performance, but it adds the

complex valving necessary for bipropellant use, and reliability past

2 years has not been proven. LOX-LH has the highest performance of

any propellant combination, but it is not storable for long periods of

time, so it is ruled out for all but the earth departure stage. Solid

fuels are storable for long periods of time and have intermediate

performance. Their main disadvantage is they must be burned to

completion. If they are selected, the AAC thrusters should be

increased to make up for this loss of flexibility. These possible

propellants are compared in Table PP-5.
POSSIBLE PROPELLANTS

39

ADVANTAGE/

PERFORMANCE

HYDRAZINE

Isp (apx) 235

Storable >12 years

complexity medium

flexibility high

average 1.008

specific gravity

Sutton p.206-9

used in yes

deep space

used for deep yes

space missions

LOX-LH N204/MMH SOLID

450 340 300

continuous 2-10 years >20 y.

losses

high high low

high high low

0.28 1.20 1.1 74

no yes yes

yes yes yes

Table PP-5. PROPELLANT COMPARISONS

Thruster Selection

The selection of the thruster type and position is discussed in

the AAC section of this report.

Fuel Selection for Each Stage

The thrusters selected to be mounted to the spacecraft will

use hydrazine propellant because of a combination of the length of

the mission (so the high reliability of hydrazine systems is
preferred), and the low delta V required (so the low performance is

not that large of a penalty). These thrusters will only be utilized in

the segment of the mission from Jupiter to Pluto.

The AAC trusters mounted on the Jupiter assist stage as well

as the main thruster for the gravity assisted delta V will utilize

bipropellant (N204/MMH). The increased performance needed for this

stage drove the decision for this propellant selection. This segment

of the mission will only last apx. 4 years, to the thrusters on this

stage will have a sufficient lifetime.

Tank sizing for Each Stage

The estimated delta V for the AAC system from Jupiter to

Pluto is apx. 1.2 Km/sec. A safety factor of 0.1 km/sec has been

added to this. This will give a total delta V for this stage of 1.3

Km/sec. From the rocket equation this gives a fuel mass of 412.5

Kg. Using the density of hydrazine, this fuel will require a spherical

tank that is 0.922 m in diameter.

The estimated delta V for the AAC system from Earth to Pluto

is apx. 0.4 Km/sec. The delta V required at Jupiter, with a safety

margin is 1.7 Km/sec, for a total delta V of 2.1 Km/sec. Again using

the rocket equation, a total propellant mass of 1075.8 Kg has been

determined. Using the densities of the fuel and oxidizer and their

mass mixture ration, the oxidizer of this stage requires a tank that

is 0.935 m in diameter, and the fuel requires a tank that is 0.956 m

in diameter.
r •

References

Schock, A. "Modular Isotopic Thermoelectric Generator"

Fairchild Space and Electronics Company. Germantown, Maryland

20767, pp. 327-342

Sutton, George P. "Rocket Propulsion Elements", 1963

Koepke, "Electric power Subsystem Requirements" AAE 241

notes, Spring 1990, Kinko's.

Koepke, "Spacecraft Propulsion System Applications" AAE 241

notes, Spring 1990, Kinko's.

41
Structures Subsystem -- Grouo 1

The structural subsystem of the PFERD mission presented an

interesting challenge. The constraints of a mission to Pluto are

formidable. The spacecraft must survive the journey while keeping

everything on the craft in working order. To this end, several

different designs were looked at

The basic structure of the craft is very important to the

craft's survivability. The nature of the mission to Pluto requires

that certain instruments are used; these in turn require that three

booms are needed. For instance, the science platform can not be

near the Radioisotope Generators (RTG's). Neither of these can be

near the magnetometer, and these should all be kept away from the

high gain antenna. This necessary configuration leads to a three

boom arrangement with the fourth component in the center.

There were three basic arrangements that were looked at, all

having a general Y-shape. The first consisted of a craft with the

science instrumentation platform at its center. The three booms

held the communication equipment, the magnetometer, and the RTG's.

This arrangement has several benefits. Since the antenna is on a

boom, the propulsion module can be attached to the center part

facing the Earth. This would eliminate the need for the craft to turn

180 ° to perform a burn, and then to turn back to reestablish

communications. This version also has its drawbacks, however. The

antenna is very large, and placing it on a boom presents stabilization

problems.
The next arrangement considered moved the antenna to the

center of the Y-shape along with an electronics bus. The three

booms hold the scientific scan platform, the RTG's, and the

magnetometer. This craft has the stability of a large structure at

its center, but it still has problems. Since the antenna is now at the

center and pointing at the Earth, the propulsion module can not be

there and is placed on the opposite side. This necessitates a 180 °

flip before a burn is performed. This presents a serious attitude

control problem, but not an unsurmountable one. The main problem

is getting the spacecraft to turn itself around without instructions

from the Earth. The next craft attempts to eliminate the need for

this extra maneuver.

The third arrangement is basically the same as the second one,

with one important change. The craft's propulsion module would be

pointing through the center of the antenna, thereby pointing in the

correct direction. This eliminates the need for an orientation

reversal, but presents numerous other problems. Since the nozzle

will be pointing directly at the Earth, it will also be pointing at the

instruments located at the focus of the antenna. These must be

protected from the hot exhaust of the engine. Also, problems in

communication due to reduction of antenna area and in sizing of the

engine need to be addressed.

Due to constraints and tradeoffs mentioned above, the second

of these arrangements was selected for use in the PFERD mission. It

is Group One's belief that the problem of turning the craft 180 ° will

be handled by existing technology in redundant computers and
artificial intelligence. Therefore, the second arrangement will be

the safest and most efficient way to reach Pluto.

A sample drawing of this craft has been included on the next

page. The observant reader will note the similarities in design

among Group One's choice and previous craft flown by NASA, such as

Voyager, Mariner, and Galileo (Dumas, p. 535). This is not just

coincidence. Not only do these designs make good sense, but they

also use many current technologies and manufactured parts that

would be easily available to the PFERD mission. This will reduce

cost and time required to complete the project, both of which are

benefits to the PFERD program.

In addition to the components mentioned above, the craft will

have micrometeorite shields outside the bus, protecting the antenna,

and protecting the RTG's. The U-shaped scan platform will provide

much of its own protection.

The spacecraft will also have a "fourth boom". The midcourse

booster will be attached to the bus on the opposite side of the high

gain antenna with explosive bolts to allow it to be jettisoned after

firing. This will provide the boost at an intermediate planet to

attain the required velocity to make it to Pluto.

The craft will have a pair of spherical fuel tanks slung under

the boom above the midcourse booster. These will provide fuel for

the attitude control thrusters and an additional maneuvering engine.

Finally, we must consider the layout of the craft before it

reaches orbit. The PFERD mission will utilize the Titan IV to
II

L.._

r_,'l

3

0 "--D, 0

/

a

"33

X'x,x," /

\

CO

,qp,

d°
PFERD CO_FIGURATION

\
q.Sc

Electronic Bus Layout

M,

1.75 m
put the probe into Earth orbit, and therefore the craft must fit in the

payload bay of Titan IV. Utilizing a folding antenna and retractable

booms, the probe will be able to fit into a cylinder 3.5 meters in

diameter and 5 meters tail, well within the range of the Titan IV's

capabilities. (see launch configuration diagram)

Once the spacecraft's general shape has been determined, the

materials used to construct it must be examined. The driving

factors in material selection are low weight, high strength, good

temperature ranges, and resistance to radiation.

High strength and low weight are the initial considerations for

the materials. Composite materials have the best strength to

weight ratio, but are very expensive and are resistant to loading in

only one direction. Titanium and Aluminum are both very strong and

very light, with Titanium being the better of the two. Aluminum is

available at a much lower cost, however.

Almost all metals are resistant to radiation, so this is not a

factor in their tradeoff studies. However, composite materials have

been known to suffer degradation due to radiation exposure in the

space environment. (AAE 241 notes, Set #9) This makes them a

poor choice for external structural components of the probe.

Temperature is also a factor in material selection. Designers

must worry about metal evaporation at high temperatures, which is

demonstrated in the following chart. The metal will lose 0.040

inches in one year at the corresponding temperatures.
_7

Table STR 1

Metal Temoerature (°C_

Cd 1 20

Mg 240

AI 81 0

Fe 1050

Ti 1250

Mo 1900

Ta 2300

MgO 1090

From Space Materials Handbook, page 498

Titanium looks to be the better choice for high temperatures.

The PFERD probe will spend most of its time in the outer solar

system where temperatures are very close to absolute zero, but the

high temperatures will be present while the probe is in the vicinity

of the earth and inner planets. However, in choosing metals for use

in spacecraft structures, one does not usually worry about the

effect of the environment because most metals have very similar

properties in space. (Space Materials, p. 640) It is only a design

using composites that must take these factors into careful account.

Due to the above constraints, Titanium was selected as the

main structural material for the booms, platforms, and structure

supporting the instruments on the high gain antenna dish. These all

will have exposure to the sun for the greatest time period. Titanium

also can be used to construct the fuel tanks. Aluminum will be
used to construct the bus and micrometeorite shields protecting the

craft on its journey. Since the craft has been modeled after several

existing spacecraft, the choice of materials and launch

configuration shown will allow the craft to withstand the loads

during launch.

The following chart presents the approximate masses of the

general components of the spacecraft. The structural masses are all

estimated from various figures found in JPL Div. 35 Mass Estimation

Reference for the Galileo and the Voyager probes modified to fit the

PFERD mission needs. The subsystem masses are from the

individuals responsible for that subsystem. At the bottom of the

chart is the final estimate for the cruise craft, excluding the

propulsion modules.
@

Table STR 2

Electronic Bus

AAC 72

CCC 5

Science 1

Thermal 1 5

Structure 61

Power 10

Antenna Structure

Antenna 30

Support 20

RTG Boom

RTG's 30

Boom 1 5

Science Platform

Boom 20

Scan Platform Struc. 26

Science Instr. 60

AAC 22

Thermal 3

Magnetometer Boom

Instrument 3

Boom 1 0

Fuel tank (empty)

fuel tank thermal control

Total Non-propulsion mass

Mass (in kg) Totals

164

50

45

131

13

46

4

-,453 kg

A complete breakdown of each subsystem is available at the

end of this section.
Thermal Analysis

In order for a spacecraft to survive the journey to Pluto and be

in working condition at the time of arrival, the temperature must be

strictly controlled. The temperature of the space environment is

near absolute zero, and instruments must be kept within a certain

temperature range. There are two ways of controlling temperature:

passive thermal controls and active thermal controls. (Space

Materials, p. 99)

Passive controls use no power or moving parts. They consist

of components such as multilayer blankets, reflective and

absorptive panels, and coatings to control the temperatures of the

interior of the spacecraft. These controls are very reliable, but are

less precise. Also, very accurate knowledge of the thermal

conditions in space is necessary to use passive controls. A passive

control also generally results in a minimum weight and minimum

cost system.

Active controls utilize power and/or moving parts to regulate

temperature. Some examples of active controls are heaters,

selective exposure disks, and fluid transport refrigeration/radiation

devices. These controls give very precise temperature control, and

they do not require accurate knowledge of the environment. As with

all mechanical devices, the potential for failure exists. (Space

Materials, p. 99)

The PFERD spacecraft will utilize a combination of these two

methods. Since the majority of the cruise will be far away from the

sun, the primary concern must be keeping the craft warm.
Since the instruments must be kept in a relatively narrow

range of temperatures, the active controls will be used more than

the passive ones. The active controls will provide continuous heat,

while the passive will keep high temperatures near the sun at bay

and improve heat retention far away from the sun.

During the cruise, the craft will be shielded from the sun by

the high gain antenna. It is this design that renders the craft

insensitive to changes in solar intensity. It is also the reason so

many craft (such as Mariner and Voyager) use a similar structural

design. (Dumas, p. 536)

The PFERD mission applies knowledge gained from previous

space missions to control temperature. Thermal control is divided

into certain areas: Bus, scan platform, and other boom control.

Bus thermal control consists of the isolation from solar

heating provided by the high gain antenna, multilayered insulation on

all sides to prevent thermal gradients, radioisotope heating units

(RHU's) strategically placed in the bus, and thermostatically

controlled louvers in the panels. (Dumas, p. 538) These measures

provide a stable thermal environment inside the bus to allow

operation of the instruments.

The PFERD scan platform will be shaped like a large three

dimensional U. This will provide thermal protection for the

instruments while allowing certain instruments to be able to view

the environment (such as the Astros 2).

The RTG's will require no heating, as they are actually

producers of waste heat. An attempt was made to harness and

utilize this excess heat, but sufficient transport media was not
discovered. Heat pipes were investigated, and it seems that an

osmotic system might have the capability to transfer the heat, but

the added weight and complexity of an osmotic pumped heat pipe did

not fit with the requirements in the Request for Proposal.(Tanzer,

p.184-5) The magnetometer will require no heating units, also. The

RHU's used will interfere with the instrument's performance.

The thermal subsystem approximate mass is presented below.

The masses are based primarily upon data from similar spacecraft,

such as the Mariner. The reader will note large differences in the

figures for the Galileo and those for the Mariner. This is probably

due to the vastly different missions the two flew. Although both

were scheduled to end up at Jupiter, the Galileo flew first inward

towards the sun, requiring more heat protection.

52

Table STR 3

Estimated Thermal Subsystem Masses

Galileo probe

Mariner Jupiter-Saturn

scheduled for 1977

PFERD mission

80 kg (JPL doc)

11 kg (Dumas, p. 542)

20 kg (see mass section)

As the reader may note, the mass of the thermal subsystem

can vary greatly according to mission plans.
53

Conclusion of Structural Section

This report presents Group One's view of the best structural

subsystem for the Pluto mission, which has been dubbed PFERD. It

has followed the Request for Proposal in its tradeoff studies and

design. The structural subsystem has optimized weight and cost as

much as possible in the choice of materials and thermal equipment.

All of the materials considered in this report exist at the present

time, so that requirement is taken care of. The Titan IV is being

utilized, and a diagram is shown of how the craft will fit inside of

the payload bay, and the Titan should easily be able to lift the entire

mass of the craft. Simplicity has been stressed throughout, and as

stated, most of the components have been flight tested on previous

missions and can be relied upon. The structural system will exceed

the mission life, because there is no practical limit on the

materials, and the RTG heaters will last easily out to Pluto ( in

excess of 20 years). Off the shelf design is being utilized across

the board, as shown by the craft being modeled after several craft

that have already flown.

In summary, the PFERD mission will deliver the "spacehorse"

(the actual craft) to Pluto safely and will return valuable scientific

data to Earth.
MASS TABLE

SU_--_S"_']_% Part

1 fine sun sensor

1 microcamputer

electronics, support

Star tracker, FUP_

control thrusters

fuel needed for cruise

1 microcomputer

1 high gain antenna

electronics

Location

bus

bus

bus

scan platform

bus

tank

bus

antenna struct.

bus

Weight (kg)

1.75

.i

i0

n/c

.I

5

platform instruments

magnetc_eter

magnetcmeter electronics

power/Prop_ ulsion

RIG's

fuel tank

electronics/cables/power

Mission Mmnaaement

scan platform

own boGm

bus

own _mm

below bus

bus/mx= 

3

1

30

i0

none
bus structure

panels

shielding

connectors

science bocm

scan plat fo_m/machinery

magnetcmeter bocm

R_G_

antenna support

bus theza_l protection

blankets

heaters (I0)

louvers and machinery

scan platfomn heaters

scan platform blankets

fuel tank heaters

fuel tank blankets

coatings

bus

bus

bus

m

antenna struct.

bus

bus

bus

I

I

m

26

5

2O

26

i0

2O

3

2

i0

1

2

2

2

1

TOTAL 453.9 (-FUEL)

r
RQfQrences

Structural Subsystem

Group 1

PFERD.

Proaress in Astronautics and Aeronautics

Vol. 35: Thermophysics and Spacecraft Thermal Control

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

Spacecraft: A Seven Mission Summary" (M.I.T.,

Cambridge, MA, 1974) : pp. 531-546.

Vol. 86: Spacecraft Thermal Control, Design, and Operation

"Osmotic Pumped Heat Pipes

(AIAA, New York; 1983): pp.

Tanzer, H.J. and Fleischman, G.L.

for Large Space Platforms"

182-201.

Soace Materials Handbook; Lockheed Aircraft Corporation.

(Lockheed Missiles and Space, Sunnyvale, CA; 1965): pp. 95-

177, 493-509, 639-640.

Jet Propulsion Laboratory Division 35; Mass Estimation Reference

(JPL, Pasadena; 1980): Data on Thermal and Structural masses

for Voyager and Galileo.

AAE 241 Class Notes, (Kinko's Copies, Champaign, IL ; 1990)

Set No. 9: Last Page, structural materials data.

Set No. 13: Inertia data; pp. 2-6.
MISSION MANAGEMENT. PLANNING AND COSTING SUBSYSTEM

57

MISSION CONFIGURATION

The first consideration was that of the type of mission to be

flown to the Pluto-Charon system. Depending upon which approach

is taken, the requirements placed upon the probe design, trajectory

type and the propulsion system vary greatly. Because of the

preliminary nature of the exploration of Plutionian space, a lander

was considered extravagent as a first mission and was ruled out

immediately. Thus, the selection for the PFERD spacecraft was a

decision between two possible configurations: orbiter and flyby.

Initially, both missions fulfilled the RFP requirements in the

scientific domain. The orbiter would provide a greater amount of

data return as opposed to the flyby mission, albeit at a greater cost

due to the increased complexity of the probe and delivery system.

Once the trajectories were examined, however, the delta-V required

for the orbiter at Pluto became prohibitively large. Thus, by default,

the configuration chosen for the PFERD mission was that of a flyby.

Upon close examination, the flyby probe could easily accomplish the

most compelling scientific objectives at Pluto, and do so at a cost

that would make it a very viable mission.
58

TRAJECTORY DETERMINATION

The trajectory for PFERD had to fulfill several requirements,

both stated and implied in the RFP. Specifically required by the RFP

was a launch date between the year 2000 and the year 2010. Other

considerations in the RFP imposed additional limitations. By stating

that components available on or before the year 1999 must be

utilized for the mission design, the launch vehicle selection was

limited and thus the range of trajectories was narrowed further.

Minimization of cost would suggest a short mission duration and a

minimum delta-V requirement as well.

In determining the optimum trajectory for PFERD, three types

of transfers were considered. The first of those was the direct

trajectory, with no intermediate encounters en route to PlutoCharon. This type of transfer has the benefit of a short flight time

and simple navigation, but requires a large delta-V at Earth.

The next two types of transfers examined both involve

gravity-assist maneuvers at planetary encounters along the way.

The first of these uses an encounter with one of the outer planets to

increase the probe's velocity and modify its flight path. By far the

most influential body is Jupiter with a gravitational constant of

1.267x108 km3/sec 2, over three times that of the next most massive

planet, Saturn. Thus, any trip to the outer planets would inheiret a

large gravity assist if an encounter with Jupiter were possible. This

is the type of trajectory that was used by Voyager I and II during

their "grand tour" of the outer solar system.
59

With the increasing size of payloads and the unavailability of a

heavy-lift vehicle in the U.S. arsenal of expendible launch vehicles, a

new type of trajectory has been determined that uses a swingby of

Venus. This type of trajectory costs less in terms of delta-V to

deliver a payload to the outer solar system where additional

gravity-assist maneuvers can be performed. These are of two types:

Venus-Earth Gravity Assist (VEGA) and Venus-Earth-Earth Gravity

Assist (VEEGA). These trajectories are currently in use by Galileo

(VEEGA) and the upcoming Cassini mission to Saturn (VEGA) [1].

The tool used to evaluate the applicibility of various

trajectories to the PFERD mission was MULIMP, a trajectory

optimizing software [2]. This program, while somewhat clumsy and

producing occaisionally conflicting results, provided a database of

various trajectories from which the final trajectory was decided

upon. Listed in Table 1 are the trajectories investigated and the

criteria they met. After consideration of the results of MULIMP, a

decision was made in favor of the Jupiter Gravity Assist (JGA)

trajectory over the only other viable candidate, a Mars-Jupiter

Gravity Assist (MJGA). This was primarily because of the fact that

a large amount of the delta-V of the latter mission had to be

executed at Jupiter (Table 2). This would entail transporting a large

amount of propellant to Jovian space and thus drastically increase

the fuel-to-payload ratio of the spacecraft.

It can be noted here that Jupiter, having a synodic period of

about 12 years and providing gravity-assist to the Voyagers in the

late1970s, could be expected to be in a position to do so again in the

first few years of the twenty first century.
60

TRAJECTORY COMPARISON

Traiectorv

Direct

NGA

MNGA

VNGA

VJNGA

VEJNGA

VJGA

VEJGA

VSGA

VESGA

JGA

JNGA

MJGA

JENGA

SNGA

JUGA

MJUGA

Low Delta-V

o

X

X

X

X

X

X

X

X

X

Criteria

Trio time <20 years

X

X

X

X

D

X

Note:

name;

Table 1

Planet names are represented by the first letter of their

GA denotes "gravity assist"
f

JUPITER GRAVITY ASSIST TRAJECTORY

Velocity delta-V

Date _ (km/sec_ (km/sec_

2000 AUG Earth 9.77 9.77

2004 MAY Jupiter 6.237 0.91

2018 MAY Pluto 8.582 0.00

17.68 Years Total DV .. 10.68

6].

MARS-JUPITER GRAVITY ASSIST TRAJECTORY

Velocity delta-V

Date _ (km/sec_

2000 MAR Earth 9.774 9.774

2007 DEC Mars 1 9.079 0.455

2010 MAR Jupiter 7.150 7.960

2019 JAN Pluto 16.581 0.00

18.85 Years Total DV = 18.189

Table 2
62

LAUNCH VEHICLE SELECTION

As previously stated, the United States variety of launch

vehicles for a mission of this scale is at present limited to two: the

Space Shuttle and the Titan expendible launch vehicle (ELV). With

the shuttle accident of 1986 and the subsequent banning of the

Centaur upper stage from future shuttle flights, the only capable

vehicle currently available is the Titan IV. With the introduction of

the solid rocket motor upgrade (SRMU) and an improvement of around

30% mass to low Earth orbit [3], the Titan IV/Centaur G' was

selected as the ideal launch vehicle for a flyby mission to Pluto.

This increase in performance will enable the Titan IV to deliver

13,600 pounds to geosynchronous orbit, and approximately xxxx

pounds to escape velocity.

If development leads to production of the shuttle-C, this could

aslo be employed as an alternative to the Titan ELV. The shuttle-C,

with its proposed payload of upwards of 80,000 pounds to low Earth

orbit, could provide for a whole new range of trajectory options.

With the possibility of greater delivered payload, trajectories such

as the Mars-Jupiter gravity assist, which require a substancial

delta-V at the intermediate encounter body, could be made

accessable.
63

SEQUENCE OF EVENTS

=

Using the Jupiter gravity assist trajectory data

MULIMP, a timeline of mission events is presented.

derived from

2000 AUG 31

2000 SEP

2000 OCT

2004 MAR

2004 MAY 19

2004 JUL

2018 MAR

2018 MAY 07

2018 JUL

Launch by Titan IV/Centaur

Nine day launch window

PFERD spacecraft extends antenna, booms

System tests/equipment checkout

Except for course correction control, PFERD

systems shut down

PFERD system power-up, begin data taking

for Jupiter flyby

PFERD at Jupiter closest approach, 3.0 radii

Engines fired to produce DV of .91 km/sec

System shutdown

System power-up

Pluto closest approach, 3.0 radii

Pluto encounter ended, extended mission

Extended mission objectives commence
COS NG

Cost estimation for a project such as this has been found to

conform to a simple algorithm. The estimate is divided up into two

major categories: the development project, and the flight project.

The development project is subdivided into two more groups, one

comprising the actual flight hardware and the other the support

functions.

Each of these can be estimated in the number of manhours

required for each subsystem as either Recurring Labor Hours (RLH),

or Development Labor Hours (DLH).

DEVELOPMENT PROJECT-FLIGHT HARDWARE

._ Mass (kg) DLH RLH

Structures 106.0 387 126.5

Thermal control 23.0 125.2 58.35

Propulsion 56.0 535.5 131.2

Attitude & Articulation 93.85 1550 725.3

Communications 20.0 634 1 55

Antenna 30.0 1394 400.7

Command & Data 0.1 71.5 17.65

RTG Power 30.0 358 242

Line-Scan Imaging 5.0 435 136.2

Vidicon Imaging 28.0 604 229

Particle/Field Inst. 8.0 314 101

Remote Sensing Inst. 22.0 380 33.85
DEVELOPMENT PROJECT-SUPPORT FUNCTIONS

DLH

System Support & Ground

Launch+30 days Ops

Imaging Data

Science Data

Management

Development

Development

Equipment 2085.6

665.8

.37

88.1

711

65

FLIGHT PROJECT

Flight Operations

Data Analysis

DLH

11014

4681

TOTAL LABOR HOURS 21,599.62

TOTAL LABOR COST(FY77) 226,796.01
66

References:

1) Diehl R., Belbruno E., Bender D., Myers M., Stetson D. "Low

Launch Energy Trajectories to the Outer Solar System via

Venus and Earth Gravity-Assist Flybys" AAS 87-419.

2) Friedlander, A.L. "(MULIMP) Multiple Impulse Trajectory

and Mass Optimization Program" Science Application, Inc

report No SAI 1-120-383-74, April 18, 1974.

3) Kolcum, E. H. "Air Force, Contractors Predict Long Life for

Heavy Lift Vehicle" Aviation Week and Space Technology,

July 17, 1989.