N91-16053

The Phoenix Pluto Probe

Group 7

?-7o

AAE 241

University of Illinois

April 24th 1990

Group Leader / Structures

Mission Management

Attitude and Articulation Control

Command, Control, and Communications

Propulsion & Power

Science & Instnmaentation

George R. Gunning

Jeff Spapperi

Jeffrey P. Wilkinson

Jim Eldred

Dennis Labij

Meredith Strinni

https://ntrs.nasa.gov/search.jsp?R=19910006740 2019-05-09T14:36:20+00:00Z
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Table of Contents

Mission Management ....................................................... 1-1

Appendix of Equations ..................................................... 1-12

References .............................................................................. 1- 13

II Structures ............................................................................... 2-1

Introduction .......................................................................... 2-1

Subsystem Interactions ................................................... 2-1

System Layout & Description ........................................ 2-2

Mass & Inertia Configuration ........................................ 2-5

Mission Constraints and Requirements ..................... 2-7

Environmental Hazards .................................................... 2-9

NEP Interactions ........................................ ,......................... 2-10

Safety Issues ......................................................................... 2-11

Thermal Control ................................................................... 2-12

Conclusion .............................................................................. 2-12

Appendix ................................................................................ 2-14

lTIScience Instrumentation .................................................. 3-1

RFP Requirements and Compliance ........................... 3-2

Imaging Science Subsystems ......................................... 3-3

Near-Infrared Mapping Spectrometer ..................... 3-5

Photopolorimeter - Radiometer .................................... 3-6

Ultraviolet Spectrometer ................................................. 3-7

Magnetometer Subsystem .............................................. 3-8

Energetic Particles Detector ............................................ 3-9

Plasma Subsystem .............................................................. 3-10

Plasma Wave Sensor ......................................................... 3-11

Dust Detector Subsystem ............................................... ._ 3-11

Micrometeoroid Detector ................................................. 3-11

Radio Science Subsystem ................................................. 3-12

Science Timeline .................................................................. 3-14

References .............................................................................. 3-15

IV Attitude and Articulation Control ................................ 4-1

AACS Requirements ........................................................... 4-1

Design Approach ................................................................. 4-3

Design Trades ....................................................................... 4-3

Initial Configurations ........................................................ 4-4

Final Design Configuration .............................................. 4-5

Control Modes ....................................................................... 4-7

AACS Hardware ................................................................... 4-7

System Integration ............................................................ 4-11
Design Problems.................................................................. 4-11

Discussion ............................................................................... 4-12

References ............................................................................

. . 4 - 13

V Power & Propulsion ........................................................... 5-1

Requirements ............................................ .............. ,............ 5-2

Method of Attack ................................................................ 5-2

System (Primary Thrusters) .......................................... 5-3

Possible Thrusters ...................................................... 5-4

Thrusters Selected ...................................................... 5-5

Propellant ....................................................................... 5-6

Ion Dynamics ................................................................ 5-8

Specifications ................................................................ 5-9

Attitude and Articulation Thrusters .......................... 5-10

Power Source ........................................................................ 5-11

Interactions with other Subsystems .......................... 5-12

Problems ................................................................................. 5-13

Equations ................................................................................ 5-14

References .............................................................................. 5-15

VI Command, Control, & Communications ...................... 6-1

Sumnmry of Abbreviations ............................................ 6-1

References .............................................................................. 6-14

Equations ................................................................................ 6-15
Mission Management

The Phoenix probe which is our design for an unmanned probe to

Pluto has an addition which was a driving force to Mission Management.

This driving force was the potential use of a Nuclear Electric Propulsion

(NEP) system. Though this will increase cost a great deal, it's use has many

far reaching effects on the space program. The NEP will not only be at least

equal in performance to this mission, but will be shown that in the future it

will be cost and performance effective for many missions to come.

Although nuclear power is under the propulsion subsystem, it has such

an effect on trajectory and other options that I must study the two,

trajectory and propulsion together, to reveal it's true merits for

interplanetary travel. The Nuclear Electric Propulsion system has many

strong points that lend themselves to the use in such a mission. The strong

points for NEP include a continuous supply of power especially away from

the sun, low acceleration, and possible trip time savings. These trip time

savings are good for long distance mission such as missions past Mars, but

are not usable for manned missions. NEP also has a low fuel consumption

and high specific impulse, thus making it attractive for missions with a high

delta-V, which is definitely a problem when going to Pluto. Another reason

NEP is attractive for the Phoenix Probe is the long life time of these reactors,

allowing long duration missions with heavy payloads. In fact their was a

study done which showed that for more missions expected of a vehicle the

cost for NEP decreased. A final point for the use of NEP is that they are safe,

increase reliability, and are operationally flexible.

With all these benefits, many of which apply to our probe, we decided

to fly an Orbiter mission. The following chart lists the reasons that an

orbiter was the best vehicle to fly.

1-1
Flyby Orbiter

Scientific: Minimum Time Sufficient Time

Cost: Inexpensive Expensive

Payload: Light Load Heavy Load

Misc: No Benefits Future Uses

Lander

Maximum Time

Very Expensive

Heavy Load

Unknown surface

As shown on this chart for a Flyby a chemical Propulsion system would be

best suited since a Flyby would not utilize a NEP systems strong points. If

we consider the distance were going for only one planet with no additional

benefits it does not seem to be a wise choice for a mission. For a Lander

mission the NEP system works well since it would be a high delta-V mission

with a heavy payload, but we don't know anything about the surface so a

lander would be a difficult task. We also considered a landers information

not equally beneficial for the increased cost, since Pluto is so far away. We

decided to Fly an Orbiter mission that would allow our scientific equipment

to take more accurate measurements. Measurements with the on board

photopolarimeter, solid state Imaging, near Infrared spectrometer, and

visible and ultraviolet spectrometer will give us a complete layout of Pluto's

thermal properties, landscape, mineralogy, and atmosphere. An Orbiter

mission also takes advantage of using the NEP system because it will be a

heavy load and an original design, and this new design will be a helpful

development for future spacecraft.

The development of a NEP system for our mission is a great advantage

for an Orbiter, but there are many missions in the future that would benefit

from this technology in cost, time, and

performance. In fact many AIAA papers (1,2,3,5,8) think that it is the

propulsion system of the future. One mission of the future that would

benefit is TAU-a mission to a thousand AU's. This mission is dependent on

1-2
NEP if it were to go 1000 AU's in 50 years, to make measurements of the

distances to the stars in our own galaxy. A Mars cargo transport mission is

also a mission that NEP severely out preforms chemical propulsion in the

time to get to Mars and payload carried. Therefore when the Mars initiative

begins they would use NEP to send the cargo ahead and have the astronauts

rendezvous with it in orbit. A trip back to Neptune using a nuclear propelled

Orbiter would take only 10 to 12 years. Using NEP system out performs

chemical system when constructing on Orbital Transfer Vehicle(OTV). When

this comparison of a NEP OTV vs. a chemical OTV was done it was shown that

after initial development, NEP was about $250 million cheaper. This

reduced cost over chemical is resulting primarily from reduced propellant

consumption and from the larger number of missions which can be

accomplished by the single nuclear stage. As shown all these missions plus

others are severely benefited by the use of NEP, therefore the sooner it is

developed, the sooner it can be implemented to these missions.

The Selection of a launch vehicle for this mission was narrowed down

by the fact that our spacecraft weights 24,914 kg. Therefore we could

initially eliminate the possibility of using most of todays U.S. launch

vehicles, with the exception of using possibly two Titan rockets. We could

use two commercial Titans, or Titan 4NUS (Type 1 or Type II). The problem

with this would be that we would have to assemble our spacecraft in orbit,

which could be done at the space station, but the cost to do all this would be

higher than launching it in one launch vehicle not to mention an on-orbit

assembly cost.

Another possible launch vehicle would be the Soviet Union's Energia.

This launch vehicle is capable of delivering payloads weighing more than

100 tons( 90,800 kg.) into a low earth orbit.(6) This payload weight should

be sufficient to lift our spacecraft to LEO, plus an upper stage, to lift it into a

nuclear safe orbit of approximately 700 km. The obvious difficulty with this

1-3
is securing the use of Energia from the Soviets. The politics of such an act in

itself would be a large accomplishment and if political breakdown occurred

then we would be stuck with an expensive spacecraft stranded on the

ground.

Other than these two options all the other worlds current launch

vehicles can be excluded from evaluation because they would need multiple

launches to get our spacecraft in orbit. The cost would be astronomical and

on-orbit assembly would be almost impossible, thus satisfying the RFP

requirement of minimizing on-orbit assembly. To make our mission at all

realistic in a cost and possibility standpoint a requirement is for the U.S. to

develop a Heavy Launch Vehicle(HLV). This development is already being

considered and planned to satisfy the future needs of NASA.(7) Studies

established that a cargo vehicle with increased lift capability (> 100,000

lbs.(-45,400 kg.)) would be required by the mid-1990's, to satisfy

anticipated civil, commercial, and defense needs.(7) The main goal in these

developments is to bring the cost of lifting vehicles to $300/1b of payload

delivered to LEO.(7)

The Shuttle-C vehicle can satisfy a variety of missions and meet

emerging payload requirements.(7) As currently envisioned the Shutfle-C

will be a launch vehicle capable of delivering a minimum of 100,000 lbs.

(45,400 kg.) of usable cargo to an altitude of 220 NM (407 Kin). The vehicle

will be operational in the late 1994 time frame and will incur minimal

facility impacts and developmental costs.(7) The Shutfle-C plus an

appropriate upper- stage should be able to get our Phoenix probe into a

Nuclear Safe orbit(NSO). Therefore the Shuttle-C is the most likely choice for

the Phoenix probe and this covers the requirement of identifying the use of a

space shuttle.

The final considerable launch vehicle would be the Advanced Launch

System(ALS). The objective of the ALS program, being jointly developed by

1-4
l,

!

DoD and NASA, is to define a launch system with a vehicle capable of placing

payloads up to 200,000 lbs. (90,800 kg in low earth orbit at a fraction of the

cost of today's launch systems.(7) This system design is being cost driven to

reduce the total delivery cost to orbit to one-tenth of the anticipated cost for

the Titan 4. In addition the launch vehicle must be highly reliable, easily

supported and maintained, and responsive to changes in mission

requirements.(7) This system has some conflicting information in that some

articles say it will be available in the late 1990's while others imply a much

longer development time, which I have a feeling is more likely. If this

system is in operation at our prescribed launch date it will definitely be the

launch system of choice by a cost standpoint.

Out of all of these vehicles the Shuttle-C will probably be our launch

vehicle. Shuttle-C is most likely to be ready on time for our mission, cheaper

than two vehicles, and easier and more dependable than using Energia, since

it will be U.S. made.

To begin in the design of a trajectory I had to first determine what

planets would be possible to flyby and thus making the design able to

preform several possible missions, an RFP requirement. To determine this I

plotted the planets in their approximate positions, at the time that our

spacecraft could reach them, with a Earth launch window between

2000-2010.(figure 1) For example Uranus is located where the dark arc is

on the circular orbit. The dates on that arc are from 2010 to 2020 assuming

an approximate nip time to that distance of ten years. This launch window

from 2000-2010 satisfies the RFP requirement. As can be see from this

figure, none of the outer planets (Saturn, Uranus,or Neptune) will be aligned

with Pluto, therefore these planets are excluded from consideration. Mars

and Jupiter are a different story, they will be lined up with Pluto during our

launch window. Mars' position is not shown on (figure 1) because it will

travel around the sun approximately five and a half times during the launch

1-5
_,'PROXIMATE POSITION s

)._,TERPLANETS AT TIME OF

,SS&Bl E-FL-"i" 6 Y'S. W [ lT _ J -.

LAUNCH D_TE BETWEEN

OF _'!i _3_.''-'x

2000- ZO__LO
window. A flyby with a gravity assist at Jupiter should give us a

tremendous acceleration out to Pluto, so I will try to include this in our

trajectory. The other possible flyby's would be Mars and an asteroid. The

Mars flyby would be beneficial to help the Mars initiative by searching for a

landing site. With reference to the asteroid it is NASA policy that all

missions that transverse the asteroid belt should include an asteroid flyby if

at all possible, which should not be to hard with 12,000 asteroids out there.

Once I considered what possible missions could be done in addition to

our Pluto Orbiter I began our trip to Pluto. Ftrst we launch the spacecraft up

into Leo and then we use an upper stage, most likely a Centaur, to put the

Spacecraft up in a NSO orbit of 700km. At this point we deploy many of the

spacecraft booms and scientific equipment. Finally we turn on our Nuclear

Electric Propulsion system and our trip begins.

The First part of this trip is to get out of Earth's sphere of

influence(SOI). The choice's are to either spiral out of the SOI or to insert

into heliocentric space with some booster. The spiral trajectory was chosen

because it has a lower mission cost and this spiral out trajectory has direct

relevance to future electric propulsion orbit transfer vehicles. The actual

spiral trajectory of our Phoenix probe looks very similar to figure 2. The

approximation I received using Cheby2 indicates it will take close to 232.3

days to spiral out to escape velocity. During the spiral away from the Earth

our spacecraft will revolve around the earth nearly 900 times, thus allowing

time for a system checkout. The velocity at NSO will be 7,452 m/s but as the

spiral continues it will slow to a final speed of 958 m/s at SOI escape. The

last 50 days of this spiral can be seen to be flattening out, this is because the

Sun's gravitational influence is becoming stronger than the Earth's. At

925,000 km. from the Earth our Phoenix probe will reach the edge of the

Earth's SOI and the origin of the system switches from the Earth to the Sun

and our interplanetary trajectory begins.

1-7
/

/

/
All low thrust trajectory analysis was accomplished using the computer

code CHEBYTOP2(Chebychev Trajectory Optimization Program). Cheby2 is a

multi-purpose trajectory program to optimize either mass or power for low

thrust trajectories of either NEP or Solar Electric Propulsion(SEP). I used it to

allow simple estimates for variable power from different planets with spiral

escapes and spiral capture. The basic information that I used includes;

Mass=20,750 kg, Isp=5500, Power=-100 kW, Propulsion system specific

mass= 57.3, and a power level of 87%. A technical problem that I had was

that most of the numbers stated within this paper are at most rough

estimates, since this program does not allow for many options and the use of

it was limited by the lack of knowledge of its internal working and proper

inputs.

The interplanetary travel begins just after leaving Earth's SOI with a

solar system speed of close to 30,500 m/s. I ran two scenarios on Cheby2.

The first one was a trip from Earth directly to Pluto. The second case

prepared consisted of a mission from Earth to Pluto with a swingby at

Jupiter. The first case from Earth directly to Pluto included a spiral out of

Earth's SOI and a spiral into a elliptical orbit around Pluto. The launch date

is to be 2451546 Julian date(JD), Jan. 3,2000, and took approximately 18.5

years. The trajectory when mapped onto galactic map does not look very

efficient, this might be caused by the fact that Cheby2 optimizes for power or

mass and not for time. This case takes a very long time and is an unlikely

choice although our probe could survive that long. This scenario requires the

propulsion system to be on for roughly 17.1 years, which our system could

handle since it has a lifetime of approximately twenty years. This trip time

is again just approximate and with some optimization for time it could be

reduced.

The second case of a trip to Pluto with an Jupiter gravity assist came

out to be more realistic. The trip time was close to 15 years, with a launch

1-9
date of 2453095 J'D, April 2, 2004, and an arrival date at Pluto of 2458599

JD, April 30, 2019. This trip time of 15 years (5504 days) is more realistic

and a better choice over case 1. While analyzing the data for this case I

noticed that the trip from Earth to Jupiter, the first 1100 days, seemed very

inefficient and has room for improvement. The propulsion system was

required to be on for roughly 14 years, thus allowing a great deal of

propulsion on time around Pluto. These numbers are just approximations

with little or no time optimization.

The reason I stress that these numbers from Cheby2 are

approximations is because out of a couple of sources(3,9) information was

given for trajectories to Neptune. These missions to Neptune are almost

exactly like ours to Pluto, because they use an Orbiter mission, Isp values of

5300 to 5978, and power of 100 kW. The only difference is the fact that

they are going to Neptune instead, but in the year that we are planning our

mission, Pluto is only 3 to 6 AU's farther away. These papers list trip times

of 10-12 years to Neptune, therefore to go an extra couple of AU's shouldn't

add more than possibly two years. This indicates a trip time to Pluto of

12-14 years.

A comparison of flight times to get an Orbiter to Pluto using chemical

propulsion is just about the same. In fact the best trip time I got with the

lowest delta-V was over 15 years also. So there are really no savings in the

way of using chemical propulsion, in fact NEP might even get us there faster

considering the mass of the Orbiter.

These missions that I planned show no Encounters with Mars nor

asteroids. These are not included because Cheby2 does not allow such

additions to your flight path. These missions would be very likely to be

included although I was unable to determine when the could occur if they

could occur. Another obstacle to find an asteroid flyby is to do this there

would be a lengthy process of going through 12,000 asteroids and finding

1 -10
those that are near our optimal trajectory.

The orbiting of Pluto is interesting in that on the way there we will

have to reverse our thrust vector to begin slow the spacecraft down so that it

can enter orbit around Pluto. This reverse thrust should begin to occur 4.6

years before Pluto is reached. Also we will have to do trajectory checks with

our sensors to define our position here and along the whole mission to stay

aligned with our trajectory. This is very important with a NEP system for we

need a longer time to correct trajectory discrepancies. The f'mal insertion

into orbit around Pluto will be a spiraling right into an elliptical orbit. With

the NEP propulsion system lasting long enough to do all of the scientific

studies of Pluto we should be able to raise our orbit and do scientific studies

of Charon. The end of our mission will occur when the NEP system finally

gives out and we receive no more communications from our Phoenix probe.

The two reactors on board should last us up to twenty years and this lifetime

is long enough for an adequate safety margin to meet the RFP requirement

of being able to carry out our mission plus others. With all this information

I have assembled a time line (figure 3) that use case 2.

Costing for our mission is done on figure 4, which itemizes the direct

labor, recurring labor hours, and total cost for each subsystem. Our mission

cost comes to $4.215 billion to complete whole mission minus the cost of the

launch vehicle, which was unattainable since the Shuttle-C is not built yet.

This cost estimate includes four spacecraft to be built, thus satisfying the RFP

requirements. Although this is an exuberant amount of money you have to

weigh this with the new cost efficient subsystem that are being designed,

especially the propulsion system. The development of the NEP system is

approximately one-third the total cost, so otherwise if this was taken out of

our costing the spacecraft would be more cost effective. This price is in

disagreement with the RFP, but again one must weigh that against the

originality of such a project and it's future benefits.

1 -11
en¢_in_ -fig'"o 4

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D

p 1 _ ,,_,,izedcostinc

' 2_

i

!31

_ _C'_ _,t'_ _IC'i _,._,.',J'rt,",.DMI_ILl M_lIT PROJECT

Mass (kg)

_trHmltr_ _ PtcIV_c_

6 ITr,erma Contro I

I

Direct L Hours

I 1000 hours

Recurring L.H

1000 hours

TotalCost

millions$

' i ",I.,58 4796 854 158.7

i 49.4 5.4

r ropu,_,,un i

8 Attitude 8, Articulator"

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II

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RTG Power

Landing Radar/Altir eter

,.'lne-_4;an Imaging

Particle & Field Inst",;ments

16 IP,emote Senslng!nstr;ments

17 5an,ptlr,g Instrumepts

'18 ' I_l II'_i"_Al'%"'_ I"=l lllf"--:,'_'_lt/C '_

I 9 <''¢_''-''" Support

20 Lounrn +30 gaysOpe'_tlons

21 i!_n._a_ng Data de,,eloz-ent

22 !S,:iencceDataDevelo?"nent

2si'- u, OO2j_]) .iiaEaoemer_:

24 _ r__:GHT PRO,_E...

,_===_

25 If'lightOperations

26 jData_.na!vs_s

_51 158

_n_ I _vuv I 12604

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Totalcostin_197 _ dollars=_')°44 o Million

io.al_ cost in l_90dollars =$4_a146Million

ORIGINAL PAGE IS

OF POOR QUALITY
Appendix of F.__uations

Cheby2 Equations

X+kx/r^3=a where X=Positi'on vector

a=Acceleration vector

k=Gravitational constant of sun

r=lXl

Constant Isp

lai=(ao/u)(p/po )o (t) & du/dt=-(ao/c) (p/po) ¢_(t)

where: ao = Initial acceleration an 1AU

c=Exhaust velocity

u--relative mass of vehicle

o (t)= 1->powered or 2->coast

Costin_ Eauations

TC=(IOO%-Z)NRC+RC

NRC=DLC-RC

DLH=DLH(2,M)+(N-2)*(RLH(2,M))/2

where: TC= Total cost

NRC--" Non-recurring cost

RC=Recurring cost

DLH=Direct labor hours

RLH=Recurring labor hours

1 -12
References

1. Galecki, D. L. and Patterson, M. J., "Nuclear Powered Mars Cargo

Transport Mission Utilizing Advanced Propulsion," AIAA Paper No. 87-1903,

June-July 1987.

2. Mahoney, S. P.,"Nuclear Powered Ion Engine Orbit Transfer Vehicle

Design and Operational Effectiveness," AIAA Paper No. 86-1391, June 1986.

3. Aston, G. ,"Ion Propulsion Technology Requirements for Planetary

Mission Applications," AIAA Paper No. 85-2000, Sep.-Oct. 1985

4. Parvez, S. and Holman, J.M.L., "Low Thrust Insertion Into Orbit Around

Mars," AIAA Paper No. 87-0093, Jan. 1987

5. Ramsthaler, J.H., "Comparison of Advanced Nuclear and Chemical

Propulsion," AIAA Paper No. 88-2927, July 1988

6. Gubanov, B.I., "Energia - New Soviet Launch Vehicle," IAF Paper No.

88-172, Oct. 1988

7. Branscome, D.R. ," NASA Launch Vehicles- The Next Twenty Years," IAF

Paper No. 88-189, Oct. 1988

8. Nock, K.T., "TAU - A Mission to a Thousand Astronomical Units," AIAA

Paper No. 87-1049, May 1987

9. Garrison, P.W. ,"Nuclear Electric Propulsion Spacecraft Configurations

study," AIAA, April 1983

10. Friedlander, A. ,"CHEBYTOP2 (Chebychev Trajectory Optimization

Program)," Science Applications International Corporation, Sponsored by

Boeing.

11. Costing Estimate done with Notes given in AAE241, Spring semester,

1990, University of Illinois, TA: Koepke, A.

1 -13
RFP

NEP

C3

STC

A&AC

HGA

LGA

I&C

MAG

I-fiR

JPL

ACRONYMS

Request For Proposal

Nuclear Electric Propulsion

Command, Control, & Communication

Structural and Thermal Control

Attitude and Articulation Control

High Gain Antenna

Low Gain Antenna

Instrumentation and control

Magnetometer

High Temperature Radiator

Jet Propulsion Lab

Introduction Structures and Thermal Analysis

The structural analyst in the Phoenix space probe serves three roles;

structural design, thermal control and material selection. It is the

responsibility of the analyst to make sure that the space probe maintains

its integrity for the entire mission. Therefore it will be shown that the

Phoenix probe meets its requirements in the Request For Proposal. (RFP).

Each requirement will be presented along with a description of how this

requirement is satisfied. A design configuration will be illustrated along

with a description of each component and its interaction with the other

components. A mass / inertia configurations will be shown as well as

descriptions of launch vehicle compatibility, on - orbit assembly, materials

selected, thermal control considerations, and safety issues of Nuclear

Electronic Propulsion (NEP). Also, a description of how the structural

analyst interacts with the science, propulsion, attitude and articulation

control, command, control, and communication (C3), and mission

management will be presented.

SUBSYSTEM INTERACTIONS

Structures and Thermal Control (STC) is a highly interactive

subsystem. STC must work with Mission planning in order to maintain low

2-1
costing, select a compatible launch vehicle, and most importantly develop a

spacecraft configuration that is ideal for a Pluto orbit insertion mission. For

the science subsystem STC must provide a clear field of view for the

scientific equipment, and maintain equipment at normal operating

temperatures. STC provides Attitude Articulation and Control (A&AC) with

approximate masses and inertias so that we will maintain stable flight. As

with science, STC must maintain C 3 equipment at ideal operation

temperatures and provide a clear field of view for the High Gain Antenna

(HGA) and Low Gain Antenna (LGA). And finally, Power and Propulsion

plays a very important part with STC. The reactors provide 100% of the

thermal control for the Phoenix. Also the highly radioactive plume and

reactor play a major role in the placement of components.

SYSTEM LAYOUT & DESCR/PTION

Numerous NEP spacecraft configurations have been proposed. Figure 1

illustrates the Phoenix Pluto probe. In this configuration the thrust vector

is orthogonal to the vehicle longitudinal axis and the reactor and payload

are at opposite ends. The side thrust and end reactor configuration was

selected because this design avoids many of the conflicting subsystem

requirements that will be discussed later. A clear field of view are

provided for the high temperature power system. Thermal control

problems are minimized by integrating the spacecraft subsystems along the

thermal gradient. 2

The power module consist of two reactors, a Reactor Instrumentation

and Control (I & C) subsystem, shield, heat transport subsystem, power

conversion subsystem and the heat rejection panels. The total length of the

deployed power module is 11.3 m with the heat rejection panels extending

to a diameter of 6.9 m. There are two attitude and articulation thruster

units attached the power conversion system directly along the z - plane.

2-2
/ POWER MODULE

x

5.8m

r = .592m

23.00 m

43.3m

figureI

2 Reactors

Reactor I & C

Shield

Heat Transport System

Power Conversion

High Temperature Radiator

(HTR) panels

a.k.a, heat rejection panels

PROPULSION

Propellant Tank

6 Main Thrusters

PAYLOAD

Main Platform (AAC housing)

Science & C3 Housing

HGA (4.8m diameter)

LGA

MAG boom (13m)

2-3
Phoenix

Launch Ready

x

12.6 m

'r

fi_u-¢ 2

The propulsion module is placed on the center of gravity to ininimize any

unwanted torque due to the thrust. Mercury propellant will be stored in a

cylindrical vessel attached directly behind the main thruster unit. The

main thruster unit will include the six thrusters needed for our mission.

Placed 23 meters down the truss is the payload module. The payload

module consist of a main structural platform with a 4.8 m diameter HGA,

LGA,Magnetometer (MAG) boom, and a science and communication housing

attached. The main platform is designed to house the four reaction wheel

assemblies used by A&AC. The science and communication housing

2-4
features four panels that are kept closed during the majority of the mission

in order to protect the equipment from contamination. Once we reach

Plutonian orbit and the thrusters are turned off, the science panels are

opened allowing a full field of view of Pluto's surface.

Figure 2 shows the Phoenix in takeoff configuration. Notice that the

High Temperature Radiator (HTR) panels fold upward. The A&AC thrusters

retract in Power Conversion module. The Power and Propulsion boom also

retracts into the Power Conversion module and the Payload boom retracts

into the payload main platform. On the payload platform the MAG boom

retracts and the HGA antenna folds up into its stowed configuration.

Completely stowed, the Phoenix has a length of 12.6 m a diameter of

3.6 m and mass of 20,914 kg (see table 2). The shuttle C is being designed

for a 4.57 m diameter, payload length of 25 m, and payload mass of 45,359

kg. Plenty of room and mass is available for packing to insure a safe

takeoff.

r

MASS AND INERTIA CONFIGURATION

A summary of the mass breakdown is shown in table 2. A

contingency of 20% of the total (dry) system mass is included. The net

payload module is 1852.6 kg. An interesting note is that an additional 570

of payload could be added without any additional cost in terms is system

interactions. This was calculated with torque and thermal gradient

considerations. As shown the net power and propulsion system dry is 5576

kg. But propellant adds an additional 12,000 kg. The subtotal (wet) came

out to be 20,914 kg. This mass is only 5.1% different from our initial

estimate made during the response to the proposal. Figure 3 shows the

simplified diagram of the Phoenix that was used to calculate the mass

moment of inertias. The values of these inertias may be found in the

appendix.

2-5
Table 1 Phoenix Subsystem Mass

ITEM DESCRIPTION

SCIENCE

IMAGE SCIENCE SUBSYSTEM

NEAR INFRARED MAPPING SPECTROMETER

INFRARED SPECTROMETER

PHOTOPOLARIMETER RADIOMETER

EXTREME ULTRAVIOLET

ULTRAVIOLET

MAGNEIXDMErER

PLASMA WAVE SENSOR

PLASMA SENSOR

COSMIC RAY

DUST DEIECIOR

HEAVY ION COUNTER

CELESTIAL MECHANICS

RADIO PROPAGATION

RADIO MAPPING

COMMAND CONTROL & COMMUNICATION

S/X BAND ASSEMBLY

ANTENNA CABLING

DATA STORAGE SYSTEM

COMMAND DE'rEL-'IDR UNIT

RFS

HGA (PARABOLOID)

LGA (HALF-WAVE DIPOLE)

UNCERTAINTY

ATHTUDE ARTICULATION & CONTROL

TWO AXIS SUN SENSOR (2)

INERTIAL MEASUREMENT UN1T

STAR SENSOR ASSEMBLY

FOUR REACTION WHEEL ASSEMBLIES

PAYLOAD MODULE STRUCTURE (INCLUDING BOOM)

POWER & PROPULSION (DRY)

PRIMARY THRUSTERS (6)

A A & C THRUSTERS (12)

REACTOR (2)

SHIELD

HEAT TRANSPORT

REACTOR I & C

POWER CONVERSION

HEAT REJECTION

POWER CC & D

STRUCTURE

SUBTOTAL, LESS CONTINGENCY

CONTINGENCY (20%)

SUBTOTAL PHOENIX (DRY)

PROPELLANT

SUBTOTAL PHOENIX (WET)

MASS(kg)

156.5

30.0

19.5

8.2

5.1

12.3

5.2

5.3

7.2

13.2

10.0

8.5

4.4

10.0

7.6

10.0

350

4.7

3.5

8.6

10.0

50.0

200.0

50.0

23.2

46.1

3.2

15.0

4.3

25.6

1300

5576

636.0

340.0

1280.0

860.0

445.0

210.0

315.0

835.0

370.0

285.0

7428.6

1485.72

8914.32

12000.0

20914.32 kg

2-6
PHOENIX MASS/INERTIA CONFIGURATION

1 2

__ (SPHERE)

-

Y

14.25 tit,,..

X

3

(CYLINDER

O

26.40

44.1

6

(aillengths areinrneters)M1=5,928kg M2= 12,7632kgM3=2223.12kg

figure3

i :

Mission Constraints and Reouirements

Here is a description of a few of the constraints and requirements

given by our project manager and implied by the structural analyst. For

concise listing table 2 illustrates the requirements related to the

design and summarizes where they are met.

According to the RFP all materials must be available up until 1999. All

structural materials for the Phoenix already exist. The support booms are

currently flight proven. And the science and communications module will

be similar to that of the Galileo and Voyager. But the thermal control of the

SP-100 propulsion system has not been thoroughly tested. According to J.F.

Mondt of the Jet Propulsion Laboratory (JPL) the generic flight system of

the SP-100 will be proven reliable by April 1995.3

The use of off-the-shelf hardware is very important in the design of

Phoenix. First of all it reduces design and development cost that should be

2-7

a

structural
Table 2

REQUIREMENT

I. Demonstrateunderstanding of RFP.

2. Describe technical approaches used to comply with

RFP.

3. Identify critical problem areas.

4. Include sensitivity analyses and tradeoff studies.

5. Describe method of attack.

6. Spacecraft must adapt to space environment.

7. Materials used available before 1999.

8. Identify & _'e on-orbit assembly.

9. S/C should have sufficient lifetime plus reasonable

safetymargin.

10. Stress reliablility, low cost, simplicity.

11. Weight & cost should be opt/m_zed.

12. S/C should be able to perform several missions.

13. Off-the-she/f hardware should be used.

14. S/C should not be a threat to environment or public

safety.

15. Show & identify layout of components & size.

16. Verify launch vehicle compatibility.

17. Give approximate mass & inertias.

18. Describe S/C thermal analysis.

19. Identify materials used.

20. Show interaction with other subsystems.

Structural And Thermal Design Requirements

COMPLIANCE

Throughout paper.

Done in each section.

Done for each section.

Done were applicable.

MISSION CONSTRAINTS&RQMNTS.

ENVIRONMENTAL HAZARDS &

NEP INTERACTIONS

MISSION CONSTRAINTS&RQMNTS.

SYSTEM LAYOUT & DESCRIPTION

MISSION CONSTRAINTS&RQMNTS.

MISSION CONSTRAINTS&RQMNTS.

MISSION CONSTRAINTS&RQMNTS.

MISSION CONSTRAINTS&RQMNTS.

MISSION CONSTRAINTS&RQMNTS.

SAFETY ISSUES

SYSTEM LAYOUT & DESCRIPTION

SYSTEM LAYOUT & DESCRIPTION

MASS/INERTIA CONFIGURATION

THERMAL ANALYSES

Done in each section.

SUBSYSTEM INTERACTIONS

directed towards the developing SP-100 propulsion system. The storable

HGA, MAG boom assemblies have been featured on the Galileo.

Unfortunately, since the Phoenix is such a unique spacecraft, most of the

structural components will have to be built for its special configuration. For

2-8
example, its 30.7 m boom assembly and payload design will be unique. But

on the other hand, the materials and methods used to construct these

components have been available and flight proven. For example, Carbon

fiber/epoxy a light weight, high strength and stiffness material with a

tailorable coefficient of thermal expansion and 15 years of proven

experience will be used in the boom assembly and support trusses. 5

ENVIRONMENTAL HAZARDS

The Phoenix Pluto probe has a complicated array of environmental

hazards that it will encounter. First is the wide range of temperatures that

exits from Earth's atmospheric temperature at take-off to Pluto's orbit that

will extend to approximately 34 au for our mission. At these distances the

temperature can reach a chilly 42 K. To protect the Phoenix from the

effects of such cold temperatures, measures must be taken to keep the all

systems within its operating temperatures. These measures will be

outlined later in the Thermal Control description.

A second environmental hazard is the meteoroid environment. Large

meteoroids are rare in space. Therefore it can be assumed for the purpose

of this mission that we do not have to design for this condition. But on the

other hand the more numerous smaller meteoroids can present a problem.

The effects of these micrometeoroids can be compared to a sandblasting

operation 1. Three systems will be in need of protection; the thin HTR

panels, support booms, and the science and communications module. To

protect the HTR panels Beryllium Armor will be exposed to the outside

surface. To keep the boom assembly from unnecessary exposer it will be

enclosed in a single layer Kapton sock. And finally the science module

shielding will be roughly equivalent to that of the Galileo spacecraft (0.5 cm

aluminum). 2

A third environmental hazard is radiation. Radiation destroys the

2-9
orderly structural arrangement of the metals used in spacecraft. Radiation

will come from two sources. The first is natural spaceradiation and the

second is the nuclear reactor and exhaust plume. Usually a NEP type

spacecraft takes longer to escapeearths radiation belts so radiation

shielding is important. But in comparison, the Galileo spacecraft was

designed for an intense Jovian environment, and the radiation exposer of

these two spacecraft are similar. 2 A detailed description of radiation

protection can be found below in the NEP Interaction description.

The final environmental hazard is spacecraft charging. As a spacecraft

becomes charged, the electrical conductivity can negatively effect the

performance of all electronic equipment.

_'EP INTERACTIONS

Basically there are two different sources of interaction with the

spacecraft by the SP-100 system. Radiation from the nuclear reactor and

effects of the propulsion system.

The SP-100 reactor produces both gamma and neutron radiation fluxes.

Therefore in order to protect immediate equipment in the HTR, a shield

must be present between the two systems. The shield is placed directly

behind the reactor and consist of both gamma and a neutron shield. The

shield is designed with tungsten as the gamma shield and beryllium as the

neutron shield. Lithium-hydride separates the two shield since the

materials are not compatible. 2,4

There are various interactions from the propulsion system that interfere

with the spacecraft; 1) surface erosion, 2) film deposition, 3) plasma

interactions, and 4) electromagnetic interference. Surfaces exposed to the

thruster beam can be eroded. Erosion can cause failure in structural

members and thermal control surfaces. The corrosive zone of the exhaust

plume is typically 15 ° but could extend to a 40 ° maximum. 2 So in order to

2-10
prevent surface erosion the thrusters point away from all components and

the HTR panels will not extend into the 40 ° cone of the thrusters. The

deposition of propellant and non-propeUant films on surfaces can cause a

serious problem. Propellant and non-propellant sputtered from the

thrusters may travel upstream due to diffusion an electromagnetic field

effects. These f'dms can alter electrical conductivity and impact antenna

performance and thermal properties. 2 The propulsion system is not in

danger of these effects because the temperature of these systems is too

high to allow these particles to condense on there surfaces. To combat

these effects, scientific equipment will be stored in the science and

communications module and instruments such as the antenna will be

blanketed for protection. The third propulsion interaction is plasma.

Plasma generation can cause spacecraft charging and arcing. Circuit logic

and breakdown of electrical insulation are results of plasma generation.

These problems can be controlled be neutralizing the beam. 2 The final

propulsion interaction, electromagnetic interference is produced by

permanent magnets and dynamic electromagnetic fields. To prevent such

interference, the thruster subsystem should be electrically isolated from

other portions of the spacecraft.

SAFETY ISSUES

One of the key requirements of the Phoenix program is safety to Earth's

population and environment. The SP-100 has been designed to remain

intact and subcritical for a wide range of accident situations, including

water immersion, flooding, burial, launch explosions, and reentry.

The unirradiated Uranium 235 fuel does not present a biological hazard.

It can be handled and worked around without any special precautions. The

reactor will remain unirradiated during ground and launch operations. The

shielding around the core prevents the reactor from going critical in the

2-11
• u

case of water flooding. And the core is honeycombed constructed with

absorber rods that protect it from blast or impact. The SP-100 has also

been designed with redundant shutdown mechanisms with two

independent control systems. To prevent damage during any possible

reentry, the nose cone of the reactor is designed with carbon/carbon

composites which have demonstrated the ability to increase its strength as

the temperature increases. One additional safety feature is that operation

of the reactor will not occur until the spacecraft has reached nuclear safe

orbit of 925 km. This orbit is high enough that radioactive elements will

decay before its eminent reentry.

THERMAL CONTROL

One of the largest problems with the SP-100 is that it dissipates so much

heat. For most spacecraft one would be concerned about keeping the

various system equipment at a temperature that is warm enough for

normal system operation. The SP-100 radiates 2.6 MWt at a radiator

temperature of 800 K. heat flux at the radiator is 23,600 W/m 2 which is

approximately 17 times the solar heating intensity. To avoid over heating

of the science and communication module, at least 21 meters must separate

the radiator and the module. (See figure 4). 6 This separation reduces the

incident heating on the spacecraft to 1400 W/m 2. To help dissipate the

heat into space a system of heatpipes and HTR panels are used. Titanium

potassium heat pipes filled with lithium fluid located in the beryllium

radiator panels accept heat directly from a source heat pipe assembly. For

a detailed description of the heat transport subsystem see fig 5.3

CONCLUSION

To sum up, the Phoenix Pluto probe will should prove to satisfy the

structural and thermal requirements described in the RFP. The over all

2-12
t

! •

_r

r

?

if'

:._. . I_ AL_AItl

G:

F,

i

0

x

1o

!

,I

- Sa_-t_ IA0OAr(_I _P_JIA_IE.

T I em._a_w

.... T ltQ_L

I

L

30 40 _0 i0 90 " I0 _1_

3_

! . |

Fig._

tOOOm32

_00.1 l

m

'10 o |,2

I.Oo t2

13Z

SpacecraFtheatingenviro_nts from SP-IO0radiators.

F_LD

jo01_0-r

III

I it

-II

_W

HEAT PtPE

TEM

I_JMF

ORIGINAL PJ_CE _g

OF POOR QUALITY

Fig.._" Heat Transport Subsy_te_ Components
configuration provides the ideal probe that is required to study Pluto. This

is exemplified by the excellent field of view that the science and

instrumentation will have. Further more, the SP -lO0 is the ideal method

of thermal control. Not only does it provide ample heat, but also much

valuable room on the payload module is saved since all thermal control

control comes from the power module.

2-13
APPENDIX

Mass Moments of inertia:

Iy - 3/10 (M1) (rl) 2 =.3(5928)(1.8) 2 =

Eauations

_2

Ix = Iz = 3/5 M1 (1/4 rl 2 + h 2) + M1 y2ffi 3/5(5928)[1/4(1.8)2 + 5.82] + (1425)2 =

2_,fE.Ckeaa 2

SPHERE

Ix = Iy = Iz = 2/5 M2 r22 = 2/5(16763.2)(.592) 2 = _2

CYLINDER

Iy - 1t2 M3 r32 = 1/2 (2223.12)(1.8) 2 = _2

Ix = Iz = 1/12 M3 (3r32 + L 2) + M3 y2 = 1/12 (2223.12)[3(1.8) 2 + 5.32] + 2223.12(26.4) =

MASS INERTIA TOTALS

Iy=

Ix =Iz=

.

6

.

.

°

°

References

Abraham, L. H., "Structural Design of Missiles and Spacecraft",

Book Company, New York, 1962.

Buden, David and Garrison, P. W., "Space Nuclear Power System and

the Design of the Nuclear Electric Propulsion OTV", AIAA Paper 84-1447, June

11 - 13, 1984.

Mondt, J.F., "Development Status of the SP -100 Power System", AIAA

Paper 8%2591, July 10-12, 1989.

Harty, R.B., "A Comparison of the SNAP Aerospace Safety Program with

Current Requirements", IECEC Paper 889244, 23rd IECEC Conference,

1988.

McCormell, V.P., "Advanced Composites Make a Case in Space",

Composites, Sept./Oct. 1989.

Zafran, S., "Conceptual Arcjet System Design

SP-100 Mission", AIAA Paper 89 - 2596, July 10-12, 1989.

McGraw-Hill

Advanced

Considerations for the

2-14
Science Instrumentation

• -

1.0 INTRODUCTION

It has been sixty years since a so called planet named Pluto has been

discovered, and scientists still do not know exactly what it is. Existing

theories state Pluto may well be a planet, but other theories argue that this

mysterious entity may be an escaped moon of Neptune or a planetesimal.

Basic quantities such as Pluto's albedo, diameter, and density are presently

unknown. Scientists believe Pluto to be composed of rock, water-ice,

methane-ice, and possibly argon. Charon, Pluto's only known satellite, is

even more mysterious than Pluto. Without the knowledge of the above

listed measurements, Pluto's and Charon's exact compositions can not be

determined.1 A spacecraft must be sent to the Plutoian system to determine

this information. The PHOENIX orbiter, equip with many scientific

instruments, is proposed to do so. Although the study of the Plutoian

system is the main objective, another goal is to obtain valuable information

about Jupiter, Mars, the asteroid belt, and any comet, asteroid, or body the

mission may encounter during its planned journey.

Twelve scientific instruments will be used during the course of the

Phoenix mission. Four are remote sensing instruments, six are fields and

particles instruments, and one is a radio science instrument. The remote

sensing instruments are of most importance to the Phoenix mission because

they will be able to unlock many of the mysteries the Plutoian system holds.

The fields and particles and radio science experiments will correlate

information of this type received by previous missions. A detailed

description of these instruments instruments are found in part 3.

3-1
2.0 REQUEST FOR PROPOSAL(RFP) REQUIREMENTS AND COMPLIANCES

1.) RFP required an unmanned probe to Pluto:

PHOENIX mission complied by developing unmanned spacecraft.

2.) RFP required mission that maximizes information while minimizes

cost:

PHOENIX complied by selecting an orbiter with a hope that a needle

probe may be developed in time.

3.) RFP requires no materials or techniques after 1999:

PHOENIX Instrumentation Subsystem(PIB) complied by using all

instruments with the exception of one which have previously been

tested, approved, and used. The one instrument being built of

existing technology, of new design, but of no breakthroughs in

technology.

4.) RFP required sufficient shelf-life to satisfy mission plus a safety

margin:

PHOENIX PIB complies with this demand.

5.) RFP requires mission to be able to perform several missions:

PHOENIX PIB complies with plans to study, Pluto, Charon, and any

other planet, asteroid, comet, satellite the path of the mission

allows.

3-2
3.0 SELECTION, JUSTIFICATION, AND POINTING REQUIREMENTS OF

COMPONENTS

3.1 IMAGING SCIENCE SUBSYSTEM (ISS)

OBJECTIVES. The imaging science subsystem is clearly the most

valuable scientific experiment carded by the PHOENIX orbiter. Scientists

believe Pluto will have a thin or nonexistent atmosphere during the

scheduled PHOENIX mission. 1 This will permit an excellent opportunity for

an accurate determination of the morphology and geology of Pluto and

Charon's surface. The ISS will also map spatial changes in color and albedo,

and monitor the variations with time. Other objectives of the ISS will be to

locate of the spin axes and rates of rotation of Pluto and Charon. The visual

images obtained by the ISS will aid in relating data acquired by other remote

sensors to certain features of the plant's surface. 3

One of the advantages the PHOENIX orbiter offers over a fly-by

mission is that the orbiter revolves around the Plutoian system allowing the

entire system to be imaged. Also, the orbiter is able to get closer to the

system's surface enabling it to take pictures of higher resolution.

When the opportunity arises, the PHOENIX orbiter will study the

atmospheres and top cloud formations of other planets such as Jupiter or

Mars. Other objects such as asteroids, satellites and comets will also be

observed when encountered. 2

INSTRUMENT. The imaging system used will be the system which is

currently being developed for the Cassini misson. The imaging science

subsystem consists of a narrow angle camera and a wide angle camera,

3-3
which share a common setof electronics.The system isbased on a 1024 by

1024 pixel charge-coupled device. The ISS iscomprised of the following

subassemblies:

FILTER WHEEL -This isa two wheel selectableoptical_tcr assembly

containing twenty-two f'flters for the narrow angle camera and fourteen

ftlters for the wide angle camera.

SHU'YFER - A dual blade, focal plane, shutter design is used. No

preparation is required before exposing an image. The shortest exposure

time is five milliseconds. There is no upper limitation.

RADIATOR - Dark current will be subdued by the passive cooling of this

radiator.

CCD - The format is 1024 by 1024 pixels, with each pixel size being 12

micrometers square. There are approximately. 50,000 electrons in the

partially inverted mode. The UV convertor lumogen phosphor.

OPTICS OF THE NARROW ANGLE CAMERA - The parameters of the

narrow angle optics are: Ritchey Chretien with three field correctors;

focal length of 2000 millimeters; focal ratio of t"/10.5; spectral range of

200-11000 nanometers; resolution per pixel of 6x6 microradians; and field of

view of 0.35 degrees square. The close-up lens in the f'tlter wheel begin to

fade out of focus at 3.8 km.

OPTICS OF THE WIDE ANGLE CAMERA - The parameters of the wide

angle optics are: refractor in type; focal length of 250 millimeters; focal ratio

of f/4.0; spectral range of 350-1100 nanometers; resolution per pixel 48x48

micro radians; and field of view of 2.8 degrees square.

Other subassemblies which will not be described here are: the detector

head, square root processor, image data compressor, director and signal chain

logic, and power supplies. For more information on these subassemblies see

reference 4.

The ISS described described above is of new design, but, will be of

3-4
existing technology. If this design is not perfected by the time of the

mission, the imaging system used on the Voyager mission shall be used

instead.

The narrow angle camera, wide angle camera, and common electronics

module will be mounted on the scan platform and inter-connected by

shielded cables. 4

3.2 NEAR - INFRARED MAPPING SPECTROMETER(NIMS)

The main objective of this experiment is to investigate the

near-infrared spectrum to determine the geology of Pluto and Charon. The

experirnent will also map and determine the mineral content of the surfaces

of these bodies.

Pluto is believed to be composed of methane-ice, water-ice, and

possibly argon, neon, and nitrogen. These molecules along with others will

be specifically monitored by the NIMS. Other objectives of this experiment

will be to probe the atmospheres and cloud layerings of Jupiter, Saturn,

Mars, and any other objects with atmospheres when the opportunities arise.

INSTRUMENT, The NIMS was selected because it combines imaging and

spectroscopic abilities in the same instrument. The telescope subassembly

consists of an all- refractive telescope with a 22.9 cm aperture Ritchey

Chretien. The focal length is 800 mm with an aperture of f/3.5.

The spectrometer subassembly consists of: a Dall-Kirkham type of

collimator, a wide angle, flat field camera, and plane grating. The collimator

has a focal length of 400 mm and a ratio of f/3.5. The camera's focal length

is 200 mm, with a f/1.75 focal ratio. The grating is dual blazed, with 400

lines per mm.

3-5
The detectors (fifteen) are of the most sensitive type available, indium

antimonide. They require cooling by a passive radiator to 80 K. Each of the

15 detectors is placed in different areas to sample specific regions of the

spectrum. The NLMS is designed to measure wavelengths in the range of

0.7 to 5 micrometers.

The NIMS consumes an average of 8 w, and weighs 18 kg. The Galileo

carried a NIMS of the above type. The NIMS will be positioned on the scan

platform near the ISS. For more information on this instrument see

reference 2.

3.3 PHOTOPOLARIMETER - RADIOMETER (PPR)

D.,B./P,,_(_I_ The primary objective of the PPR experiment is to measures

the polarization and intensity in the region of visible fight (400-700

angstroms). This data will yield information about the properties of

fight-scattering surfaces. 3

A second objective will be to measure the thermal radiation of Pluto

and Charon. Another objective is to find the radiation budget of the Plutoian

system by measuring the total thermal emission and reflective solar

radiation. 2 The above stated objectives will also be applied to the

atmospheres of Jupiter and any other planet with an atmosphere when

encountered.

INSTRUMENT. The PPR used on the Galileo mission was the instrument

selected to be carried by the PHOENIX mission. It was selected because of

its dual abilities to measure photometry and infrared radiometry. The

instrument is equipped with a Dall-Kirkham telescope with 10 cm aperture

and a 50 cm focal point. This is the primary optical path of the subsystem.

This optical path collects light and passes it through selected f'dters. This

collected fight is then measured by detectors.

3-6
There are two minor optical paths in the PPR. The first of these paths

gathers radiation from the surveyed object. The other minor path collects

radiation from space. These minor optical paths are used only in the

radiometry mode of the instrument. Infrared channels in the radiometry

mode are set below 4 micrometers, at 17, 21, 27.5 and 37 micrometer, and

above 42 micrometers.

In the photopolarimerty mode, only radiation entering the primary

optical path is emitted to the detectors. A beam is passed through a filter

and enters in to a Wollaston prism. By rotating the filter wheel, the

polarization of the transmitted beam rotated 90 degrees. This determine the

orientation of the polarization of the incident beam. Polarimetry channels

are centered at 4100, 6780, and 9450 angstroms. Photometry channels are

centered at six positions between 6180-8920 angstroms.

The PPR subsystem has three important safety features: deployable

covers which shield all optical when thrusters are fired, sunshades which

prevent sunlight from directly entering, and replacement heaters which

maintain the temperature when the power is turned off. The PPR subsystem

weighs 4.8 kg, uses a peak power of 10 watts, and is mounted on the scan

platform with the other remote sensing instruments. 2

3.4 ULTRAVIOLET SPECTROMETER(UVS)

OBJECTIVE. The main objective of this experiment is to determine the

structure and composition of the atmospheres of Pluto (if there is one),

Charon, and any other satellite of Pluto which may exist. Atmospheric gases

discharge radiation at ultraviolet wavelengths for two reasons. They are

sometimes excited by bombardment with energetic particles, and sometimes

the resonance dispersion of solar ultraviolet radiation cause this. 3

Airglow will be analyzed by the UVS. The UVS will also determine

3-7
t"-

¢.

l •

L •

i

ultraviolet reflective properties of the surfaces of these bodies. This will

yield information to help characterize surface materials and their physical

state. 2

INSTRUMENT. The PHOENIX mission selected an ultraviolet

spectrometer similar to the instrument carded by Galileo. This instrument

consists of a Cassegrain telescope (250 mm aperture), a monochromater,

three detectors (photomultipliers), and control logic. The telescope is unique

in that it can sample ultraviolet radiation coming from a small portion of the

atmosphere or surface. The field of view produced by the spectrometer is

0.1 by 1.4 degrees for 1100-1900 and 2800-4300 angstrom detectors and

0.1 by 0.4 degrees for the 1600-3000 angstrom detector. The

monochromator has a focal length of 125 mm.

A programmable grating drive which is regulated by the control logic

controls the wavelength of the radiation being measured. The grating

supplies a resolution of 13 angstroms in the first order spectrum and 7

angstroms in the second order spectrum. The photomultipliers are capable

of investigating wavelengths from 1150-4300 angstroms. Photon pulses are

counted every 0.0007 seconds. This UVS was selected because of its wide

range of spectra (1150-4300 A °) and its flexibility in variety of data taking

programs .2

The UVS subsystem weighs 5.21 kg, and consumes 5.33 W at 2.4 kHz

and 50 Vac. It is secured on the scan platform with the previous three

instruments .2

* NOTE: No direct sunlight can enter any of the remote sensing

instruments.

the sun.

All instruments shall be equip with shields to block

3.5 MAGNETOMETER SUBSYSTEM(MAS)

OBJECTIVE, Interplanetary space is traveled by the solar wind, streams

3-8
k

of charged particles, and shifting magnetic fields that the solar winds bring

with them. Some planets have their own magnetic fields. The main

objective of this experiment is to determine if Pluto and Charon possess

magnetic fields. The second objective is to investigate interactions between

Pluto's and Charon's magnetospheres, if any exist.

The magnetometer experiment will also acquire data on all other

magnetic fields encountered during the Phoenix mission. This data will be

used in comparative studies with data received from other fields and

particles instruments.

INSTRUMENT. The magnetometer subsystem consists of four

subassemblies; two high field magnetometers (I-IF), which measure + 0.5G

to + 20G, and two low field magnetometers (LFM), which measure + 8.8

gamma to + 50,000 gamma. The Phoenix orbiter does not spin, therefore the

type of magnetometer that was carried on the Voyager mission will be used.

Each of the four subassemblies consist of triaxle fluxgate

magnetometers that measure field and intensity along three orthogonal axes

simultaneously; thus, producing direct vector measurements. One LFM is

placed at the middle of the boom (0.80 kg), and the other is placed at the

end(0.75 kg). This arrangement will allow the spacecraft's magnetic field to

be separated from the ambient magnetic field. In doing this, accurate

information can be obtained. Both I-IFMs are placed near eachother, at the

proximal end of the boom (0.26 kg each). The total mass of the MAS is 5.72

kg. 3

3.6 ENERGETIC PARTICLES DETECTOR(EPD)

OBJECTIVE. The main objective of this experiment is to investigate the

temporal fluctuations and spatial disbursement of ions and electrons in the

3-9
medium to high energy range (0.015 to 0.2 MeV and 0.1 to 1.0 MeV

respectively). This experiment will be performed in the Plutoian system,

interplanetary space, and other systems when encountered.

INSTRUMENT. The EPD has two bidirectional detector telescopes which

are mounted on a platform in the spun instrument section. The telescopes

used are a low-energy magnetosphere measuring system (LEMMS) and a

composition measuring system (CMS). The LEMMS includes an ion telescope,

two detectors, and a magnetic electron spectrometer. The energies measured

by this subassembly are .015 to 0.2 MeV and 0.1 to 1.0 MeV. The CMS is

comprised of a three-parameter detector system consisting of nine detectors.

These detectors measure the energy spectra, composition, and pitch angle

distributions of energetic ions in the Plutoian system. The EPD subsystem

has a total mass of 10.77 kg and is located on the spun instrument section. 2

3.7 PLASMA SUBSYSTEM(PLS)

Plasma is gas found in space that is electrically neutral, but,

composed of charged particles. The main objective of the PLS experiment is

to measure plasmas velocity, density, and pressure. PLS instrument also

determines the plasma flow direction by measuring the variation velocity

with r fion.

INSTRUMENT. The PLS subsystem used on Galileo was selected over

the PLS subsystem used on the Voyager for the following reasons. First, it

has an extended energy range of 1.2-50,400v; where as the Voyager PLS had

a range of 10-5920v. Second, it has three miniature mass spectrometers

which analyze ion compositions, while Voyager had none. Finally, while

Voyager's PLS had a temporal resolution of 100 seconds, Galileo's PLS has a

temporal resolution of 5 seconds. 2,3

3-10
3.8 PLASMA WAVE SENSOR(PWS)

The objective of this instrument is to identify and analyze

the radio and plasma waves in Pluto's magnetosphere. The PWS is equipt

with the the capability of remote sensing of source location. Magnetospheres

of other planets and satellites will be studied when opportunities arise.

INSTRUMENT. The PWS consists of an electric dipole antenna for the

detection of electric fields and two coil magnetic antennas for the detection

of magnetic fields. These subassemblies measure spectral characteristics of

electric and magnetic fields in the range of 5 Hz to 5.65 MHz. The total mass

of the PWS is 7.22 kg. The antennas are located at the end of the

magnetometer boom on the vertical axis. 3

3.9 DUST DETECTOR SUBSYSTEM(DDS)

The dust detector experiment will aid in the understanding

of physical and dynamic properties of small dust particle in the Plutoian

system. This information will help answer questions about the existence of

Charon, which is thought by some to be a fragmented piece of Pluto.

INSTRUMENT. The DDS is comprised of a set of grids that sense the

impacts of dust particles. The instruments field of view is 140 degrees. It

can measure masses in the range of 10 -19 to 10 -9 kg and velocities in the

range of 2 to 50 km. The DDS measures 0.1 by 0.1 m, weighs 4.37 kg and is

placed on the spun instrument section to determine the flight direction of the

particles. 2

3.10 MICROMETEOROID DETECTOR(MMD)

_. Micrometeoroids are particles smaller than one mm in

3-11
diameter that are present in the space occupied by our solar system.

Although the Voyager mission took no particular notice to the asteroid belt

due to the results of the Pioneer 10 and 1 l(no concentration within the belt),

the Phoenix mission will carry a micrometeoroid detector (MMD) to study the

belt and verify Pioneer's irmdings.

A second reason for employing this instrument is to study the Plutoian

region for these particles. A knowledge of the micrometeoroids present in

this area may unlock some of the mystery of the being of Charon. It may

give some clues as to if Charon is a fragmented piece of Pluto.

INSTRUMENT, The MMD used on the Phoenix mission is similar to the

instrument used on the Mariner-Mars spacecraft. A crystal acoustical

transducer is fastened to aluminum plates (22 cm by 22 cm). The crystal

will discharge an electrical pulse whenever a micrometeoroid strikes the

plate. The plate is completely covered with an insulting and conducting film.

This forms a capacitor sort of detector. A potential is placed across this

capacitor and an electrical discharge occurs when a micrometeoroid

perforates the insulation of capacitor. This type of capacitor detector is self

repairing and is excellent for repeated use. When the capacitor detector

output coincides with the output of the acoustic detector, the direction of the

micrometeoroid can be determined. The present design of the MMD allows

for the determination of the number and penetration power of the

micrometeoroids.

New MMDs which will calculate velocity as well as momentum may be

available before the Phoenix is built. This advanced instrument will be used

in place of the above described MMD if so. 5

3.11 RADIO SCIENCE SUBSYSTEM(RSS)

3-12
_]_./F,,__]]._'_'_'_ Two experiments, celestial mechanics and radio

propagation, will be investigated by the radio subsystem. The celestial

mechanics experiment will be used to determine the structures and shapes

of the gravitational fields of Pluto and Charon. This subsystem uses the radio

system to perceive gravitational perturbations on its trajectory.

A primary goal of the radio propagation experiment is to study

ionospheres, atmospheres, and magnetospheres. This will provide

measurements of density, pressure, and temperature as a function of height;

which is dependent on the doppler shift. While not as important for the

probing of Pluto, the experiment will be more essential for the studies of

planets with atmospheres.

INSTRUMENT. The radio frequency subsystem is used in combination

with receivers and transmitters based on earth. The RFS measures doppler

shifts, echo time delays, amplitude, spectrum and polarization of radio

signals. The mass, size, and location of this assembly can be located in the

Command, Control, and Communications subsystem. 2

3-13
4.0 SCIENCE TIMELINE

1 23

EARTH

4

1 - Instruments conduct observations of earth and Moon to calibrate

2 - Cruise mode

3 - Fields and particles instruments begin operating continuously

4 - Study asteroid belt (2.2 to 3.5 AU)

5 - Perform trajectory maneuvers to cancel launch injection errors and

refine aiming

6 - Remote sensing of Plutoian system

7 - Scan and photograph star field

* - Every 0.25 AU scanning instruments will perform remote sensing

* - Observe all planets, satellites, meteors, and comets trajectory

passes

I •

t

3-14
References

1.) CrosweU, Ken; "Pluto: Enigma on the Edge of the Solar System",

Astronomy; July,1986; pg. 8-22.

2.) National Aeronautics And Space Administration, Washington D. C.;

Galileo: Exploration of Jupiter's System; NASA SP 479, pg. 119-141;

1985.

3.) NASA Jet Propulsion Laboratory, California Institute of Technology;

"Voyager Neptune / Interstellar Mission: Flight Science Office

Science and Mission Systems Handbook"; JPL D-498, PD 618-128, pg.

1-52; October 1, 1987.

4.) NASA Jet Propulsion Laboratory, California Institute of Technology;

"Cassini Mission: Saturn Orbiter Proposal Information Package,

Volume VI: Imaging Science Subsystem Description"; JPL D-6464, PD

699-11, Vol. VI; October 10, 1989.

5.) NASA, Auburn Alabama; "JOVE: Jupiter Orbiting Vehicle for

Exploration. Volume 1", pg. 3.1-3.47; NSR 01-003-025,

CR-61880; August, 1967.

3-15
ATTITUDE AND ARTICULATION CONTROL SYSTEM

The goals of the attitudeand articulationcontrolsystem (AACS) are to

achieve and maintain a particular orientation in space. The first phase of

this process is attitude acquisition which employs a variety of sensors to

locate the spacecraft in space relative to some inertial reference frame.

Stabilization of the craft in this orientation is maintained through the use

of control actuators which must also be capable of maneuvering the

spacecraft from one attitude to another. 4 The selection of the AACS

methods and hardware depend on the mission requirements, with special

cam taken to insure compatibility and integration with the other

subsystems.

ARCS RE0 MENT 

Table 1 outlines the specified and derived requirements pertaining to

the AACS, and provides a reference location of compliance for each

requirement. The primary requirements of the AACS are to survive the

long life of the mission and be capable of several different missions. The

first of these leads to the derived requirement of total redundancy of all

systems, while the mission flexibility requirement calls for a reliable

system of control actuation. Also the fifteen-plus year life of the mission

dictates the need for autonomous control. An increasing communication

delay time as the spacecraft moves further away from earth and periods

of no communications require an on-board system capable of analyzing

attitude acquisition information and implementing control actuation to

maintain spacecraft stabilization without the benefit of command. This is

accomplished with advanced software on-board with preprogrammed

actuation sequences to accommodate all

4-1
AACS REQUIREMENTS

7"_ b le £

SFECIFIED REOUIREMENTS (R.F2)

1. Optimize performance, weight, and costs in design trades.

2. Design must be reliable, low cost, simple, and easy to

operate.

3. Use "off the shelf" hardware developed before 1999,

when available.

5. System should have a sufficeint lifetime plus a safety

maT_

6. Must be an original and imaginative design.

7. Identify the design approach and technical probl eros.

8. Probe. must be capable of several missions.

DERIVEDREO $

1. Maintain antenna and science instzun_nt pointing

2. Select a stabilization method.

3. Select types and placement of se:x_orsand actuatom

4. Integrate the AACS with other subsystems.

5. Determine torque and momentum requirements.

6. Must have partial autonomous control capability.

7. Determine environmental effects.

8. Must have a fifteen year minimum lifetime.

9. Total redundancy of all systems

4-2
conceivable maneuvering scenarios.

Further requirements of the AACS are dictated by the basic structural

configuration of the flight vehicle. For instance the dumbbell type

configuration selected for the final design must be Three-axis stabilized.

Spinning the vehicle about the pitch axis (x-axis) or the thrust vector

(z-axis) would result in poor communication capability since the antenna

must be placed at the far end of the spacecraft to avoid adverse

interaction with the nuclear propulsion system. Spinning about the roll

axis (y-axis) would result in an unstable spin which would eventually lead

to an undesirable end over end rotation about the pitch axis (x-axis). All

other requirements are dependent upon AACS component selection and

placement and are discussed throughout the report.

ii

!

DESIGN APPROACH

The method of attack for selecting the AACS is basically a design by

design approach. Following a considerable amount of initial research,

several spacecraft and AACS configurations are selected with input from

the other subsystem analysts. These preliminary design choices are then

analyzed to determine if they satisfy the real and implied requirements of

the mission. All problems with the selected systems are then outlined

and further research is done to determine possible solutions to these

problems. F'maUy the options are compared and a final configuration is

selected. The remaining analysis consists of refining the best choice and

presenting the final design.

DESIGN TRADES

The first design trades considered are low cost versus reliability, long

life, and accuracy. This cost pertains to both weight and monetary cost

and is a factor in the selection of the AACS hardware. Another important

4-3
trade related to hardware selection is an original design versus "off the

shelf' hardware. Newer components may be technically superior but

previously space tested hardware has the overwhelming advantage of

known performance parameters, which reflects the use of tested

components in the final design configuration. Other trades relative to the

final design include maneuverability versus disturbance sensitivity and

reaction wheel versus thruster control in terms of stabilization capability

and fuel consumption.

INITIAL CONFIGURATIONS

Three different spacecraft and AACS configurations were selected for

the preliminary design analysis.

They include:

1. Spin stabilized spacecraft - Chemical propulsion, RTG power.

2. Spin stabilized spacecraft - Nuclear electric powered upper

stage.

3. Three-axis stabilized spacecraft - Nuclear electric

propulsion, two on-board reactors.

The first choice is a Pioneer type scientific probe with hardware

modifications made to fulfill the mission requirements, such as long life.

This configuration was rejected without further research due to its

incompatibility with the nuclear electric propulsion (NEP) system selected

by the the design team.

The second spacecraft configuration utilizes Three-axis stabilization

throughout the initial thrust phase of the mission, which is limited by the

assumed ten year life of the NEP system. At this time the entire NEP

system is jettisoned and a spin stabilized scientific probe continues on to

Pluto powered by RTG's. The advantage of this particular configuration is

4-4
thatthe NEP upper stagecan delivera largerpayload Through the initial

delta-vrequired than a weight comparable chemical upper stage.1 Also,

following the NEP system detachment, the scientificprobe would only

requirea fiveyear power activelifetime,assuming a fifteenyear mission.

Disadvantages of thisselectioninclude a largelaunch mass and a lossin

simplicityof design. This configurationwould would requiretwo

indelx;ndentcontrolsystems, one forthe three-axiscontrolof the primary

vehicle and another for the spin stabilizedcraft.Also a largechange in

the mass of the vehicle following the NEP system detachment would

require a complex control scheme to maintain stability. These drawbacks

and the resulting high monetary cost of such a mission do not satisfy the

specified mission requirements.

FINAL DESIGN CONFIGURATION

The third preliminary configuration was selected as the final design

on the basis of mission requirement compatibility and a favorable

analysis of the design trades. A layout of the spacecraft including

locations of the AACS components is shown in figure 1. The vehicle

consists of two nuclear reactors, a power conditioning unit, and heat

shielding at one end, and the scientific payload and C^3 hardware at the

opposite end. The spherical fuel tank is located directly below the main

thruster block, both of which are positioned at the vehicle center of mass.

As discussed earlier, three-axis stabilization is the only viable control

method for this dumbbell type configuration due to the requirements of

maintaining adequate communication capability while avoiding adverse

interaction with the NEP system. Furthermore, a flexible system utilizing

active control is desirable to counteract the effects of structural vibrations

within the 28.5 meter extendible boom. 5

4-5
• ,j v,

AACS

THRUSTERS (12)

/

Y

4-6

AACS control

cylinder

(Gyros, wheels)

2-axis

sun sensor (2)

Star sensor

assembly
The three-axisactivecontrolsystem offersthe advantage of inertial

stabilization with the potential for high pointing accuracy. It is the best

method for maneuvering which allows for high precision and adaptability

to perform several different missions. A disadvantage of the system is

that six possible control directions ( pitch,roll, and yaw) must be

maintained. Also a two-axis sun sensor is required due to the absence of

rotation.

CONTROL MODES

The controlmodes for the variousphases of the mission are:

I. Attitudeacquisitionmode

2. Cruise mode

3. Trim maneuver mode

4. Orbit insertionmode

5. Large maneuver mode

The f'trst three modes rely primarily on sensor information and low

maneuvering thrust, while the last two Require both sensor information

and considerable auxiliary propulsion. Further analysis of the control

modes is discussed in terms of AACS hardware selection and performance

in the following section.

AACS HARDWARE

To fulfill the requirements of long mission life, pointing accuracy, and

total redundancy a dual control actuation system was selected. The

system includes twelve .005 newton thrust mercury ion thrusters (4 on

each axis with 6 in operation and 6 redundan 0, and a reaction wheel

assembly. Figure 2 shows an operating schematic of the system. During

the initial attitude acquisition phase of the mission both systems will be

4-7
YAW

(X)NIROL

PITCH / ROI.I.

CONTROL

Y-AXIS

X-AXIS

I

YAW

PITCH

SKE-'W

4-8
used to increase the spacecraft maneuverability. Throughout the 12-14

year cruise phase The reaction wheel assembly will provide primary

control actuation, with the thrusters used for momentum desaturization

and trim maneuvers. The final stage of the mission requires fine pointing

of the science instrumentation and the antenna, which is control by the

more stable reaction wheel assembly. Again the thrusters could extend

maneuverability or take over prirnary actuation if necessary. This

configuration satisfies the reliability requirement through total

redundancy, and minimizes the auxiliary propulsion fuel usage during the

cruise phase while maximizing maneuvering capability throughout the

mission.

The attitude acquisition system includes a pair of two-axis sun

sensors mounted on either side of the payload platform, which provides a

4I'I steradian view. A celestial sensor assembly utilizing six detector slits

in a spoke configuration is mounted at the far end of the payload platform

to allow an unobstructed field of view for continuous star reference. Also,

an inertial measurement unit containing three rate integrating gyros (2

for three redundancy) is located in the AACS cylinder centered along the

y-axis of the spacecraft, which provides displacement information

through rate integration to the control computer. 2 Figure 4 shows the

location of the attitude acquisition system on the payload platform and

table 2 describes the AACS components and gives the total AACS mass.

All selected hardware has been space tested, particularly in the

Defense Meteorological Satellite Program Which satisfies the "off the shelf"

requirement. 2 Also the system is capable of switching attitude

acquisition responsibilities to different sensor configurations in the event

of a component malfunction, which provides for total system redundancy.

4-9
Two-axis

sun sensor

star sensor

Y

Inertial measurement

unit

Two-axis

sun sensor

Figure 4

Sensor placement

Q

AACS HARDWARE DESCRIPTION

Table 2

Inertialmeasurement unit:

(Honeywell) 56.0W I_.0Kg

2. Sun Sensor(2)

(SAG_HCMM) 3.0W 1.6Kg

3. StarSensorAssembly

(Honeywell) 1.5W 4.3Kg

4. Reaction Wheel Assembly (4)

(RCA AED) 16.0 W 25.6 Kg

5. Mercury Ion Thrusters (12) ---Included in propulsion subsystem ....

TOTALS 76.5 W 46.5 Kg

4-10
SYSTEM INTEGRATION

A primary requirement of the AACS is integration with the other

subsystems. The science and communication subsystems both rely

heavily on the the AACS for antenna and instrument pointing. Antenna

pointing accuracy must be in the range of .5 to 10.0 degrees, while

instrument pointing requires an accuracy range of .35 to 2.0 degrees. The

three-axis stabilized design meets the requirements with a pointing

capability of .001 to 1.0 degrees depending on selection of and condition

of the sensors. Sun shielding is another important concern of the science

subsystem during the early phase of the mission. The initial solution to

this problem was to orient the spacecraft such that the antenna would

shield the instruments, but this approach was rejected in favor of of

enclosing the sun sensitive instruments in a hinged shield box when not in

use. Finally the configuration must be such that the center of mass does

not change as fuel is expended. To avoid this problem the spherical fuel

tank is located directly on the y-component of the vehicle center of mass.

i.

DESIGN PROBLEMS

External and internal torques on the spacecraft can cause undesirable

structural stresses and changes in attitude if not counteracted. The threeaxis active control system is particularly sensitive to environmental

disturbances such as meteoroid bombardment and solar radiation. Also

impingement forces from the ion plume effects and internal torques due

to actuator operation tend to take the spacecraft out of a stable

configuration. The spacecraft will oppose these disturbance forces with

occasional trim maneuvers to return the vehicle to the desired orientation.

Another problem imposed by the long mission life is gyro drift. To

correct this deviation the star sensor is used to obtain an exact position

4-11
k

from the last best position estimate from the gyro. This correction is

returned to the gyro and actuation is implemented if necessary. 4 Other

problems encountered include the required life of the AACS components,

which is satisfied by total system redundancy, and mercury

contamination of the sensor surfaces from the main thrusters, which is

minimized as the distance between these areas increases. The more

sensitive instruments require shielding which is accomplished with the

enclosed science box and small shields above (towards the propulsion

section) the star and sun sensors.

The final design selection meets all of the specified and implied AACS

requirements, and should provide an excellent attitude acquisition and

maneuvering system for a mission of this type. The mission is limited

only by the lifetime of the system hardware, which should increase in the

future. The AACS is particularly effective for spacecraft maneuverability

which is necessary to fulfiU several different missions. Future research

should focus on improved autonomous control capability, the radiation

effects on C^3 and science systems, and long life reactors capable of

powering a spacecraft for ten or twenty plus years.

4-12
.

Q

t

o

.

Deninger, W.D., "Electric Propulsion for Constellation Deployment and

Spacecraft Maneuvering", AIAA Paper 88-2833, July 11-13, 1988.

Cheng, GJ., "The Mars Observer: Attitude and Articulation control

system", AAS Paper 88-001, Sept., 1988.

Ninomiya, Keiken and Inoue,Hajime, "In-Orbit Control of Astro-C

Satellite (1987-012A) Attitude Control System", AAS Paper 88-051,

Sept., 1988.

Wertz, James R., "Spacecraft Attitude and Articulation Control",

Reidel Puublishing Company, 1988.

Zafran, S., "Conceptual Arcjet System Design Considerations for the

SP-100 Mission", AIAA Paper 89-2596, July 10-12, 1989.

!

4-13
PROPULSION AND POWER

A.) REQUIREMENTS

B.) METHOD OF ATTACK

C.) SYSTEM

1) PRIMARY THRUSTERS

a.) POSSIBLE THRUSTERS

b.) THRUSTER SELECTED

c.) PROPELLANT

d.) ION DYNAMICS

e.) SPECIFICATIONS

2.) ATTITUDE AND ARTICULATION THRUSTERS

3.) POWER SOURCE

4.) INTERACTIONS WITH OTHER SUBSYSTEMS

D.) PROBLEMS

E.) EQUATIONS

F.) REFERENCES

5-1
A.) REQUIREMENTS

1.) Should use off the shelf hardware, nothing which has not been

developed by 1999.

2.) Must be ready before 2010.

3.) Should optimize performance, weight, and cost.

4.) Should be reliable and easy to operate.

5.) Must be able to withstand any environment it may encounter.

6.) Must have a design lifetime to carry out its mission plus a reasonable

safety factor.

7.) Nothing in the design should preclude it from performing several

possible missions.

8.) Design will stress simplicity, reliability, and low cost.

9.) Exceptions to proposed technical requirements should be identified and

justified.

10.) Primary thrusters must be able to deliver to Plutoian Orbit.

12.) Propulsion 'and Power subsystems must not interfere with other

subsystems.

13.) Power subsystem must be able to deliver the power required by all

other subsystems at any given moment.

B.) METHOD OF ATTACK

The general process that I followed when I was designing the

various components of the propulsion and power subsystem is what I call

my method of attack. The first thing was to develop a fundamental

5-2
understanding of various types of ossibilities for a given component.

Next, I evaluated the pros and cons of each candidate for that component,

and how they related to the needs and objectives of the mission. By

process of elimination, I then determined which candidates may be

realizable. Next, I investigated the realizable candidates in depth, and

determined which one is most suitable for the given mission. Finally, I

continued to develop, and address problems related to the candidate

decided upon until the final design is complete.

C.) SYSTEM

1.) PRIMARY THRUSTERS

In determining the type of primary thrusters, several factors were

considered. First, the system should make efficient use of its propellant.

The common measure of propellant efficiency is specific impulse(Isp)

which is defined as the ratio of thrust to mass flow rate of propellant.

Thrusters with high values of Isp have high exhaust velocities which

translates to a high amount of energy in there exhaust streams. This

allows such thrusters to move a more massive payload with less

propellant. The second factor is thrust. Systems with higher values of

thrust will be able to make journeys in less time for a given type of

trajectory, either low thrust or impulsive. In addition, systems with high

enough thrust to use impulsive velocity change trajectories have the

benefit that their trajectories are computationally much simpler than low

thrust trajectories. The third factor is the ease and cost of producing the

system. In expensive systems which have been or can easily be developed

5-3
and tested are preferred. The forth factor is additional mass associated

with the system. Though a system may use its propellant efficiently, the

associated mass may make the system as a whole inefficient compared

with other options.

a.) POSSIBLE THRUSTERS

The first type of thruster considered is the chemical rocket. Solid

chemical rockets have high thrust, but low Isp. In addition they cannot be

throttled. Certain liquid propellants have an adequately high Isp to be

used as a primary thruster on a journey of this length. However, the mass

of the payload would be limited. Both solid and liquid chemical rockets

have the benefit that they have already been developed, and flight proven

many times.

The second type of thruster is the electrically propelled rocket.

This includes electrothermal, electrostatic, and electromagnetic

thrusters. These types of thrusters are capable of attaining very high

values of Isp, but generally have low values of thrust. One drawback to

this type of propulsion is that it has not really been researched on an

interplanetary scale. Another drawback is that electric methods of

propulsion require large amounts of power. This power requirement has an

associated mass which is large with respect to the rest of the system.

The third type of thruster is the nuclear rocket. Performance of

nuclear rockets is limited by the fact that there is a limit on the

maximum solid surface temperature that the reactor must operate within

to ensure structural integrity. Thus, unlike the condition found in a

5-4
chemical rocket where the energy release is within the propellant, the

propellant temperature in nuclear rockets is restricted to being less than

the wall temperatures, and hence less than that found within chemical

rockets. Another drawback is that since the propellant passes directly

through the core of the reactor, the exhaust stream is contaminated.

Nuclear rockets also have additional associated mass penalties which

come from the reactor.

The fourth type of thruster type is cold gas. This is simply the

thermodynamic expansion of a cold gas. Cold gas thrusters have low

values of Isp, but are reliable and have been flight proven many times.

Other types of thrusters are solar, and laser. Solar propulsion is

ineffective at the great distances from the sun that will be characteristic

of this mission. Laser thrusters, as of yet are not developed.(ref. 1,2,3,5)

b.)THRUSTERSELECTED

Upon evaluating the options, I decided to use an electrostatic thruster on

the Phoenix probe. During 1980, Studies at the Jet Propulsion Lab focused

on the application of nuclear electric propulsion(NEP) to outer planet

missions. The study concluded that NEP was much better than other

competitive technologies, and that a 100 kw(electric) system

significantly out performed chemical propulsion systems for outer planet

exploration.(ref. 2)

Since NEP has not been developed, In reality, It would be the case

that many additional dollars would have to be spent on research,

development and testing for this mission. This would make it very

5-5
unappealing as the best best method for the mission. However, as stated

in class by teaching assistant Andy Koepke, for this project, it may be

assumed that the technology has already been developed, and that costs

affiliated with research and development may be neglected.

The additional mass associated with the power system needed

makes the benefits of this type of propulsion system unclear when the

payload mass is small compared to the power system mass. In fact it is

possible that the propellant mass for the Phoenix probe may even be

higher than that for analogous chemically propelled missions. However

the the real benefits of NEP comes from the fact that once the mass of the

power subsystem is fixed, the marginal or additional amount of propellant

required for a given marginal payload mass will be much less than that for

a chemically propelled system. Since the RFP states that the system

should be capable of performing several types of missions, it is very

important that the system should have a capacity for a marginal payload.

Also with the capability of taking greater payload masses to a destination

also comes the capacity for designing better science experiments which

would not be realizable with chemical propulsion. Though at its present

status the phoenix mission may not appear to be the best choice in terms

of money, it has the capacity of having added to it some very advanced

science experiments, including possibly a lander, before its launch date.

In addition, information gained on NEP from this mission will be very

beneficial to future high energy deep space missions where propellant

efficiency is crucial.

c.) PROPELLANT

Determination of propellant is based on several factors. First, the

5-6
propellant should have a high nuclear mass, and a low ionization potential.

This is because the beam thrust is proportional to the square of the mass

to charge ratio. Second, the propellant should be easily stored. This is

especially important on missions of comparable duration to that of the

Phoenix mission. Third, the propellant should be environmentally safe, non

corrosive, and have minimal effects on other subsystems. Fourth, the

propellant should yield a high thruster efficiency. (ref. 1,2,5)

One possible propellant is cesium. Cesium has a high mass to charge

ratio, but is highly corrosive. Thus, it would be hazardous to both the

environment as well as the other subsystems. Another possible propellant

is xenon. Xenon is environmentally safe, and easily stored. However, it is

expensive and rare. In fact there may not be enough currently available to

make this one trip. Though xenon is a prime candidate for earth orbital

transfers, there is simply not enough to make it practical for missions

comparable in length to the Phoenix Mission. Another inert gas which

could be used is argon. Argon is also environmentally safe, but is difficult

to store. In addition, argon is more abundant than xenon. The final

propellant considered was mercury. Mercury yields the highest thruster

efficiency of those propellants considered. In addition, it is easily stored.

The main problem with mercury is that it is poisonous. Since only a small

fraction of the mission will be spent near the earth's atmosphere,

environmental contamination is not a big problem.(ref. 2) This coupled

with the fact that it best satisfies the guidelines used to evaluate the

various propellants, makes mercury the propellant selected.

The sizing of the propellant tank was done by starting with the

assumed value for the total mass of the mercury required which is about

12,000 kg. Next, the density of mercury was obtained, and turned out to

be 13,800 kg/m^3. The volume required to contain the mercury was then

5-7
computed by dividing mass by density. This gave a propellant volume of

0.87 cubic meters. Since a sphere is structurally more sound than a cube,

the propellant will be contained in spherical tank of radius 0.592 meters.

C.) ION DYNAMICS

The method which will be used to generate ions will be electron

bombardment. The neutral mercury or plasma, will be passed through a

cylindrical anode. Surrounding the cylindrical anode will be a solenoidal

coil which will be used to generate an induced magnetic field in the

direction of the plasma flow. At the center of the cylindrical anode will

be a heated filament cathode which will be the source of electrons. The

filament will be heated by passing an electrical current through it. As a

result, the heated filament will bleed of electrons. The free electrons

will be accelerated radially outward by the cylindrical anode. The

presence of the magnetic field will give a tangential force acting on the

electrons making them spiral outward toward the anode, increasing the

likelihood of them hitting a mercury atom before they reach the anode.

The collision between the electron and the neutral mercury atom will

produce the ion.

Once the ions are produced, they will then be subjected to an

electrostatic potential difference. They will be accelerated toward an

electrode which is at a lower potential. When the ions reach the

accelerating electrode, they will be at their minimum potential, and have

their maximum kinetic energy. As their momentum carries them past the

electrode, they will be accelerated back towards that electrode, and will

begin to lose their kinetic energy. Therefore it is necessary to recombine

the ion stream with an electron stream in order for the ions to retain

5-8
their momentum. Ideally one would want to recombine the ion stream

with electrons at the point of lowest potential. However, trying to do so

will result in the electrons diffusing into the acceleration field. Thus,

there is an optimal distance from the electrode that the electron stream

should be recombined with the ion stream. I, however am unable to

compute this optimal distance. The electron stream used to neutralize the

ion stream will be produced by the same method as the one in the ion

source, using a heated cathode filament.(ref. 1)

e.) SPECIFICATIONS

Since a thruster comparable to those which will be used on the

Phoenix probe has never been built, it is difficult to say how one would

perform, most of these results were obtained from tables, or from crude

approximations from similar data calculated by the Jet Propulsion

Laboratory. The information has been combined from several sources, and

in some instances represents the state of the art system which may not

be attainable.

(ref. 2)

AVERAGE THRUST ......................................................................... 0.5 NEWTONS

SYSTEM THRUST ............................................................................ 2.0 NEWTONS

SPECIFIC IMPULSE ................................................................ 5000 SECONDS

5-9
BEAM DIAMETER .......................................................................... 30 ca

THRUSTER MFETIME ...................................................... 125,000 HOURS

POWER REQUiRED/THRUSTER ................................................ 20 KWIE )

NUMBER OF THRUSTERS .............................................................. 6

NUMBER OF OPERATIONAL THRUSTERS ................................ 4

MA_RUSTER ..................................................................... 106 KG

DRY SYSTEM MASS ................................................................... 636 KG

PROPELLANT MASS ............................................................. 12000 KG

WET SYSTEM MASS ............................................................. 12636 KG

3.) ATTITUDE AND ARTICULATION THRUSTERS

The thrusters which will be used for controlling the attitude and

articulation of the spacecraft, like the primary thrusters, will be Ion

rockets. They will be very similar to the primary thrusters conceptually,

but will be on a smaller scale. In order to control the attitude of the

spacecraft, six thrust vectors will be needed. For each direction two

thrusters will be present. This makes a total of 12 AA thrusters, 6

operational, and 6 for redundancy.

(ref. 3)

AVERAGE THRUST ......................................................................... 0.005 NEWTONS

SPECIFIC IMPULSE ................................................................ 2650 SECONDS

5-10
BEAM DIAMETER ............................................................................. 8 CM

THRUSTER LIFETIME ........................................................ >15000 HOURS

POWER REQUIRED/THRUSTER ................................................... 0.2 KW(E)

NUMBER OF THRUSTERS ........................................................... 12

MASS/THRUSTER ........................................................................ 28 KG

SYSTEM MASS ............................................................................ 340 KG

!-

4.) POWER SOURCE

It is clear from the specifications for the ion rocket that a great

deal of electrical power will be required. Specifically, to run the four

thrusters will require 80 kwe. In addition, power must be reserved for

other subsystems onboard Phoenix. Development of such a power source

has been pursued intensely in recent years. The main product of this

research and development is the sp-100 nuclear reactor. The sp-100 has

an electrical power output of 100 kw. This will fulfill the 80 kw required

by the four operational thrusters, and leave 20 kw for other subsystems.

The other subsystems should not require nearly that much power. The

reactor lifetime is about 7 years at maximum power output, and longer

for output less than maximum. Since the sp-100 onboard the phoenix

spacecraft will be operating at about 82%, it will be assumed that the

reactor lifetime is 10 years. Since the mission is expected to take about

15 years, it will be necessary to bring two reactors. Another benefit of

using NEP is that it allows the other subsystems as much as 100 kw for

several years after arrival at the destination. Thus science projects

requiring large amounts of power can be conducted over long periods of

time.

5-11
(ref. 4)

THERMAL POWER OUTPUT ....................................................... 1.4 MW(T)

ELECTRICAL POWER OUTPUT ............................................ 100 KW(E)

REACTOR LIFE AT MAXIMUM OUTPUT .................................. 7 YEARS

REACTOR LIFE AT 82 % .......................................................... 10 YEARS

REACTOR MASS ....................................................................... 640 KG

SHIELD MASS ........................................................................... 860 KG

HEAT TRANSPORT MASS ..................................................... 445 KG

REACTOR I & C MASS ............................................................ 210 KG

POWER CONVERSION MASS ................................................. 315 KG

HEAT REJECTION MASS ........................................................ 835 KG

POWER CC&D MASS ................................................................ 370 KG

STRUCTURE MASS ................................................................... 265 KG

SYSTEM MASS ......................................................................... 4600 KG

5.) INTERACTIONS WITH OTHER SUBSYSTEMS

Z

J

!_

In addition to the thermal and plume interactions which are

associated with chemical propulsion spacecraft, there are also reactor

neutron and gamma fluxes as well as electromagnetic fields associated

with an electric propulsion spacecraft. Thermal interactions are

minimized by the fact that the spacecraft subsystems are integrated

along a thermal gradient. The high temperature reactor at one end,

intermediate temperature equipment in the middle, and low temperature

science instrumentation at the other end. Other interactions, as well as

thermal, are reduced by putting distance between the interactive

5-12
elements.(ref. 2) Since I do not really have an understanding of most of

these interactions, details on the configurations required by two

interactive elements was obtained from examples done by the Jet

Propulsion Laboratory.

D.) PROBLEMS

Many problems have come up during the design of Phoenix and its

propulsion system. One problem is the political pressure of having a

nuclear reactor onboard

a space vehicle. It will be difficult to convince the public that the reactor

will remain safe in the event of an accident at launch even though it has

been verified

to remain safe in almost any type of disaster.

demonstrating the true effectiveness of NEP.

Another problem has been

Almost everything in the

design of a space mission is geared tO the optimal level of Chemically

propelled rockets. When NEP performs at this level, it appears to be an

inferior method of propulsion. Thus, in order to sell the Phoenix program

it may be necessary to turn it up a notch in mission objectives as to

utilize the full potential of NEP. I have encountered many problems in the

design of the Phoenix propulsion system. Some of these problems are that

details related to this type of propulsion are difficult to find if they even

exist, and often data conflicts depending on the source. Another design

problem is that optimizing computation dealing with many aspects of the

design are difficult, or at least exceed my level of education. Thus, I am

often required to go on blind faith as to the validity of some of the

results.

5-13
/\

REFERENCES

1.)

1968

Jahn, Robert G., "Physics of Electric Propulsion", Mc Graw-Hill,

2.) Garrison, Philip W., "Nuclear Electric Propulsion Spacecraft

Configuration Study", AIAA, April 1983

' 3.) Poeschel,

i,

Technologies

q.

Robert L., "A Comparison of Electric Propulsion

for Orbit Transfer", AAIA, April 1983

4.) Mondt, J. F., "Development Status of the SP-IO0 Power System",

IAA paper 89-2591, July 1989

5.) Oats, Gordon C., "Aerothermodynamics of Gas Turbine and Rocket

Propulsion", AIAA, 1984]

.4

7  ̧

i

!

5-15

PRECEDING PAGE BLANK NOT FILMED
SUMMARY OF ABBREVIATIONS

C 3

AI

CDS

R.F.P.

D.O.D.

RH32

GVSC

COMMAND, CONTROL, AND COMMUNICATION

ARTIFICIAL INTELLIGENCE

COMMAND DATA SYSTEM

REQUEST FOR PROPOSAL

DEPARTMENT OF DEFENSE

RADIATION HARDENED 32-BIT PROCESSOR

GENERAL PURPOSE VERY HIGH SPEED INTEGRATED

CIRCUIT SPACEBORNE COMPUTER

HGA

LGA

HPBW

RTI

DSN

N.E.P.

HIGH-GAIN ANTENNA

LOW-GAIN ANTENNA

HALF-POWER BEAMWlDTH

REAL-TIME INTERRUPT

DEEP SPACE NETWORK

NUCLEAR-ELECTRIC PROPULSION

6-1
Command, Control, and Communications

Group 7

The document "Request for Proposal for an Unmanned probe to Pluto"

lists requirements which must be understood and complied with if this

preliminary design work is to be useful in the ongoing design process which will

result in the eventual construction of an unmanned probe to be sent to the

celestial body known as Pluto. While all requirements listed in the R.F.P.

(Request for Proposal) pertain to the development of the C 3 (Command, Control,

and Communication) subsystem, only those requirements which most directly

apply to the C 3 subsystem are explicitly discussed in this portion of this

document. A table listing requirements that are of particular importance is

shown below (table C 31).

TABLE C 31 • REAL AND IMPLIED REQUIREMENTS

-Select microprocessors and peripherals for Phoenix

-Select software to optimize spacecraft autonomy

-Select and size communications hardware for mission

that allows transmission at adequate speed with high

quality

-Develop overall communications plan, including ground

communications

-Recognize and defend against pointing problems and

communications loss

-Optimize mass, size, strength, reliability, cost, and

performance

-Components must be space qualified

-Provide sufficient computer speed and storage to

implement Artificial Intelligence

-Provide sufficient data storage for scientific objectives

-Utilize components available no later than 1999

-Design hardware to be redundant when possible

-Design software to be as robust and autonomous as

possible

-Transmit and receive command, telemetry, tracking and

science data

6-2
To comply with the requirements in the R.F.P. a modified

design-by-design approach was followed. Reference materials pertaining

to the C 3 subsystem were found without excluding references that did

not specifically pertain to the exact R.F.P. requirements. These

references were used to gain a general knowledge of the C 3 subsystem on

past and proposed space missions. The general knowledge from these

sources was then used to interpret the design requirements that were

imposed by the R.F.P. and by the evolving designs of the other Phoenix

subsystems. This synthesis of general knowledge, R.F.P. requirements,

evolving Phoenix probe design, and information attained from AAE 241

class notes shaped further research and design work as it applied to the

C 3 subsystem. After an initial design was reached, the subsystems were

consciously integrated and an iterative process was begun to optimize the

overall performance of the Phoenix.

A major responsibility of the C 3 subsystem design team is to select

computer equipment to be used on the Phoenix. Driving factors in the

selection of the computer equipment for the Phoenix probe were

dominated by the desire for greater autonomy than previously attempted

in spacecraft design. This desire for autonomy, specifically through the

implementation of AI (Artificial Intelligence), requires that the computer

system for the Phoenix must be faster and have more memory than past

NASA interplanetary probes. Therefore, it is important that the fastest

microprocessors available be selected and combined with a large amount

of internal memory and external storage. Three microprocessors were

seriously considered for use in the development of the Phoenix computer

system. They include the D.O.D. (Department of Defense) developed RH32

(Radiation Hardened 32-bit Processor), the Department of Energy's Sandia

6-3
Application 3300, and the D.O.D. developed GVSC (General Purpose Very

High Speed Integrated Circuit Spaceborne Computer). The RH32 was

selected due to the high speed of its 32 bit architecture and the added

reliability its radiation hardening will afford in the environment of our

Nuclear-Electric Propulsion system and the environment of Venus or

Jupiter in the event of a gravity assist fly-by. The entire computer

system will be loosely based on the multiply redundant CDS (Command

and Data subsystem) used on the recent Galileo space probe. Six RH32

microprocessors in combination with eight memory units have been

selected to be linked by a bus running at approximately 400 KHz with a

RTI (Real Time Interrupt) running at approximately 15 Hz (a

configuration similar to what was used as a part of Galileo). The internal

memory can be backed up to and loaded from an external storage

system utilizing the space proven magnetic tape that NASA has used on

numerous past interplanetary missions.

This computer hardware will be used to implement an artificially

intelligent autonomous system that has been referred to as an

"intelligent associate".l. The capabilities of an AI system, which are

expected to be available by the time of the Phoenix mission, will make

the mission more productive and versatile than it could be without the

use of AI technology. With an approximate round trip light time to

Plutonian space in the neighborhood of eleven hours, the Phoenix must

be able to carry out its mission without constant supervision from earth.

The time that it takes for a signal to be sent to the Phoenix, demonstrates

the correcmess of the R.F.P. requirement that the spacecraft design

should maximize autonomy and use AI wherever possible. Advantages

gained by the implementation of autonomous systems in spacecraft

design include a reduction of mission operation costs, an increase in

overall mission productivity, and an increase in mission success

6-4
probability. Continuing work in the field of AI will provide many

possible capabilities with which the Phoenix could be equipped.

Capabilities which will be useful and practical for implementation in the

Phoenix Probes CDS include distributed control of multiple subsystems,

fault prediction and analysis, automated real time planning and

replanning, and a reasoning/learning supervision of on-board systems.

Using sets of "heuristic algorithms" and priorities the Phoenix Probes

on-board computer systems will independently react to the changing

environments that the craft will encounter. Through an integration of

science data, engineering data, tracking, telemetry, and its programming,

the Phoenix probe will respond to threatening situations and unique

opportunities for scientific observation. The reprogrammable nature of

current spacecraft computer components will also allow mission

designers at earth a great deal of flexibility after the Phoenix has been

launched. The R.F.P. states that the design of the spacecraft should not

preclude its use for other missions, and the ability to reprogram the

Phoenix computers is an important way in which this requirement is met.

Much as the Voyager mission planners were able to send "patches" to

deal with Voyager performance anomalies, so to will the Phoenix and

Phoenix mission planners be able to respond to changing mission

circumstances and requirements. The inclusion of eight memory units

(more than twice the memory of Galileo) allows much more flexible

control of on-board systems during different phases of the mission.

When the program for a certain mission operation is no longer needed it

can be backed up to magnetic tape or discarded altogether leaving room

for new programs to be implemented in system memory. In the event

that multiple hardware failures should occur, defeating redundant design

considerations, the situation could be handled through the use of

programming "patches" which could account for the new spacecraft

6-5
performance characteristics. The extreme length of the light time from

the Phoenix to earth during most of this mission also suggests the use of a

"store and forward" command system.2. In a "store and forward" system

large blocks of commands are sent as a single communication to be

received and verified before the execution of commands is begun, as seen

in fig. C31.

FIG C31 • STORE AND FORWARD COMMUNICATION SYSTEM

MISSION OPERATIONS

CENTER

MISSION NON-REAL-TIME

COMMAND GENERATION I

STORAGE BEFORE ENTRY INTO

REAL-TIME SYSTEM [

•

t

TORAGE BEFORE FORWARDING

DSN

DEEP SPACE

NETWORK

_ORAGE BEFORE

RADIATION

SPACECRAFT

SPACECRAFT

STORAGE BEFORE

EXECUTION

EXECUTION COMMAND

It should be noted that the use of an autonomous system and the "store

and forward" technique need not preclude the use of near-real-time

commanding of the Phoenix probe. A large amount of memory also

allows redundancy in the gathering of scientific data for transmission to

the earth. Copies of images or science data can be saved in memory or

backed up to magnetic tape until confwmation of the reception of the

data can be beamed back from earth, preventing the loss of important

data taken during "one chance" scientific observations. It may also be

6-6
noted that the choice of N.E.P. and an orbiter mission will greatly reduce

the number of these"one chance" observations. It is necessary that the

C3 subsystem interact closely with all other on-board systems. The

programs implemented as part of the CDS must be able to coordinate the

activities of the power and propulsion subsystem; the attitude,

articulation, and control subsystem; the thermal control system, and the

science instrumentation subsystem. It is the responsibility of the

on-board computer to transmit its commands and commands from earth

to each of the other spacecraft subsystems.

It is also the responsibility of the C 3 design team to select and or

design the components that will be used to communicate between the

spacecraft and the earth. To accomplish this different communication

systems were considered, including laser and traditional multi-frequency

radio communication. Though technology for laser communications is

developing quickly, the desire to use off-the-shelf components when

possible suggested that the use of S and X-band communications with the

earth would be most cost effective. Often in Spacecraft communication

system design antenna gain and power required for communications

must be painstakingly evaluated to find the ideal balance between

communications performance and spacecraft mass. On the Phoenix probe

the abundant power provided by the XP-100 reactor and the overall

large mass of the spacecraft imposed new parameters to be evaluated in

the choice of spacecraft antenna. The most important factor driving the

size of the Phoenix probe antenna is the transmission data rate that will

be required to beam the science data gathered by Phoenix back to earth.

Antenna's from past NASA missions were examined to see if they might

meet the communication needs of the Phoenix spacecraft as they

interacted with its larger power system. Pointing difficulties for different

portions of the mission suggested that multiple antennas might be

6-7
p- included for use during different phases of the trip to Plutonian space.

Interaction with the structures subsystem dictated that launch volume of

the main HGA (high-gain antenna) could be minimized by using a folding

system similar to that used on the Galileo mission. A comparison of

different antenna types with respect to gain and pointing factors (HPBW,

Half-power beamwidth) was made.

This information can be seen in table C32. 3.

TABLE C32 • Antenna Type Comparison

(_9nfiguration

Isotropic radiator

Infinitesimal dipole or

loop

Half-wave dipole

Paraboloid

Gain above isotrooic radiator HPBW, deg

- v

1,0 360.0

1.5 89.9

1.64 78.0

6.3 to 8.8 (Area/wavelength*2) 60 to 70(wavelength//

diameter)

The Galileo main parabolic HGA was chosen to be used as a part of the

Phoenix with some minor redesign. It was estimated to be large enough

to meet the data rate transmission requirements of the Phoenix probes

science subsystem while still remaining small and light enough to be

launched with the rest of the craft. The redesign would involve the use

of lighter structural materials and antenna shielding, as the Phoenix HGA

will not be used as a solar shield as it was on the Galileo mission. The

Phoenix variant of the Galileo main antenna will fold to be stowed at

launch as did its predicesor. The Phoenix HGA will communicate with the

earth and DSN (Deep Space Network) using both X and S-band

frequencies. The maximum power transmitted will be approximately 1

KW. This unprecedented amount of power is a result of the unusual

6-8
fnature of our nuclear power source. The deployed diameter of the

antenna will be approximately 4.8 meters, so that a minimum amount of

redesign will be required on the Galileo antenna while still fulfilling all

the antenna requirements for the Phoenix probe. In addition to the

parabolic HGA a smaller LGA (low-gain antenna) will be used as part of

the Phoenix design. The 1 meter LGA will be a half-wave dipole antenna.

The modest increase in antenna gain over an isotropic radiator is made

up by the 78 degree pattern through which communication with earth

can be maintained using the LGA. The ease with which the Phoenix probe

could reattain contact with the earth in the event of some problem makes

this secondary antenna an important tool for increasing the mission

success probability. The LGA will also play an important role in the early

phases of the mission when propulsion concerns may be more crucial

than the pointing of instruments and the HGA. The large HPBW of the

Phoenix LGA will allow the spacecraft to almost constantly transmit and

receive engineering, tracking, telemetry, and command transmissions

should they be necessary. Fig.C32 shows a representation of the Phoenix

Communication subsystem. 4.

FIG C32 : PHOENIX COMMUNICATION SYSTEM

6-9
PHOENIXCOMMUNICATION SYSTEM

SCIENCE DATA v

_ FLIGHT DATA SUBSYSTEM ---_ TELEMEIRY.__.._

ENGINEERING DATA _ S AND X-BAND ANTENNA

GROUND BASED COMMUNICATION SYSTEM

.___.TELEMETRY DEMODULATION

MISSION COMPUTATION

CONTROL

DiN

DOPPLHRMEASUREMENT

RANGING TRANSMISSION

AND DETECTION

COMMAND PROCESSOR

A unique and important consideration in the design of the Phoenix

probe's communication system was the presence of the SP-100 nuclear

reactor and mercury ion thrusters as part of the main propulsion unit.

Though research into the effects of ion thrusters on a communication

system of this type show that impact is slight ( approximately a .2 K

increase in antenna noise temperature) 5., the general configuration of the

Phoenix probe allows the communication system to be isolated from both

the thrusters and the reactor by the main structural boom.

The design of the C 3 subsystem involved making many compromises

between the performance of a given piece of equipment and other factors

imposed by the R.F.P. and the interactions between the C 3 subsystem

and others. The speed and storage capability of the computer system

was maximized to allow for as complete as possible implementation of AI.

e

The decisions regarding command procedures were driven by a need to

make the Phoenix probe as autonomous as possible. Communication

6-10
system choices were mainly dictated by the vast distances and amount of

science data that Phoenix will beam to earth from its position orbiting

Pluto. Major design problems that have been identified include the

uncertainty about the conditions of Plutonian space, the interaction

between the N.E.P. system and communications, the relatively long life

required for this mission, and the great distance between the earth and

Pluto.

The following page shows a graphic depicting the Phoenix HGA.

The next page shows a breakdown of the major component masses of the C 3

subsystem.

6-11
PHOENIX HGA

Structural member

Feed Horn Assembly

r

Main Reflector

material

4.8 meter_

X-band Assembly

S-Band Assembly

Parabolic Reflector

Mounting Hardware

6-12
COMMAND, CONTROL,ANDCOMMUNICATION

MASSESTIMATES

S/XBANDANTENNAASSEMBLY

ANTF.NNACABLING

DATASTORAGE

COMMANDDATASUBSYSTEM

MODULATION/DEMODULATION

RADIOFREQUENCYSUBSYSTEM

MAINHGA

HALF-WAVEDIPOLELGA

5KG

4KG

19KG

35KG

10KG

50KG

250KG

50KG

TOTALMASS APPROX.423KG

COMMAND, CONTROL, ANDCOMMUNICATION

6-13
ij. -

References Cited

1. Technology for Future NASA Missions: Civil Space Technology

AIAA/NASA Conference

2. Deep Space Telecommunications Systems Engineering 1983

Joseph H. Yuen

3. Deep Space Telecommunications Systems Engineering 1983

Joseph H. Yuen

4. Deep Space Telecommunications Systems Engineering 1983

Joseph H. Yuen

5. Spacecraft Ion Beam Noise Effects

G.L. Anenberg

6-14
APPENDIX A: EQUATIONS

POWER RECEIVED

PR = PT + LT + GT + LS + GR + LR IN DECIBELS

PARABOLIC ANTENNA GAIN

G = 10 LOG10( .55 (3.14 DIAMETER/WAVELENGTH) 2

SHANNON'S LAW

B = W LOG2(PR/PN + 1) = INFORMATION CAPACITY

6-15