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N91-16049

AAE 241

SPRING 1990

FINAL REPORT

OPTIC

Orbiting Plutonian Topographic Image Craft

Proposal For An Unmanned MissionTo Pluto

APRIL 24, 1990

https://ntrs.nasa.gov/search.jsp?R=19910006736 2019-05-09T14:35:41+00:00Z
TABLE OF CONTENTS

INTRODUCTION ...................................................................................1

OPTIC DESIGN GROUP ............................................................................................. 2

1.0 SCIENCE INSTRUMENTATION ................................................................ 3

INTRODUCTION ............................................................................................ 3

1.1 DIRECTING FACTORS .................................................................................. 3

SCIENCE DIRECTIVES ................................................................................... 4

1.2 SCIENCE INSTRUMENTS ............................................................................ 6

CHOSEN SYSTEMS ......................................................................................... 6

OTHER INSTRUMENTS CONSIDERED ................................................... 8

1.3 INSTRUMENT DESCRIPTIONS .................................................................. 9

IMAGING SYSTEM ......................................................................................... 9

IMAGING PROCESS ....................................................................................... 12

MAGNETOMETER ......................................................................................... 13

ULTRAVIOLET SPECTROMETER .............................................................. 15

RADAR MAPPER. ........................................................................................... 17

APPENDIX 1.1 ............................................................................................................... 19

1.5 BIBLIOGRAPHY ............................................................................................... 21

2.0 MISSION MANAGEMENT, PLANNING AND COSTING .................. 22

INTRODUCTION ............................................................................................ 22

2.1 METHOD OF ATTACK .................................................................................. 22

DELTA-V REQUIREMENTS ......................................................................... 26

TIMEIXNE .......................................................................................................... 27

2.2 COSTING ........................................................................................................... 27

APPENDIX 2.1 ............................................................................................................... 28

FOR.MULAS USED FOR COSTING ............................................................. 28

TRAJECTORY PLANNING ........................................................................... 30

2.3 BIBLIOGRAPHY ............................................................................................... 31

3.0 POWER AND PROPULSION ........................................................................ 32

INTRODUCTION ............................................................................................ 32

3.1 POWER SUBSYSTEM ..................................................................................... 32

3.2 PROPULSION SUBSYSTEM ......................................................................... 38

3.3 CONCLUSION .................................................................................................. 46

APPENDIX 3.1 ............................................................................................................... 48

3.4 BIBLIOGRAPHY ............................................................................................... 51

4.0 ATTITUDE AND ARTICULATION CONTROL ...................................... 52

INTRODUCTION ............................................................................................ 52

4.1 SYSTEM OBJECTIVES .................................................................................... 52

4.2 ATTITUDE DETERMINATION ................................................................... 52

MODES OF OPERATION ............................................................................... 56

THREE AXIS HYDRAZINE JET CONTROL SYSTEM ............................. 58

4.2 ARTICUlaTION CONTROL ........................................................................ 61

POSSIBLE PROBLEMS .................................................................................... 62

4.4 AACS MASS SUMMARY ............................................................................. 63
APPENDIX 4.1 ............................................................................................................... 64

4.5 BIBLIOGRAPHY ............................................................................................... 65

5.0 STRUCTURES .................................................................................................. 66

INTRODUCTION ............................................................................................ 66

5.1 STRUCTU'RES SUBSYSTEM OVERVIEW ..................................................... 66

5.2 STRUCTURAL CONSIDERATIONS .......................................................... 67

MASS REDUCTION ........................................................................................ 67

STRUCTURAL LAYOUT ............................................................................... 68

MICROMETEORITE PROTECTION ............................................................ 73

THERMAL CONTROL ................................................................................... 75

ELECTRONICS PLACEMENT ....................................................................... 75

RADIATION EFFECTS ................................................................................... 75

MASS AND INERTIA CONFIGURATION ............................................... 77

5.3 BIBLIOGRAPHY ............................................................................................... 80

6.0 COMMAND, CONTROL AND COMMUNICATION ............................. 81

6.1 SYSTEM OBJECTIVES .................................................................................... 81

6.2 RFP REQUIREMENTS .................................................................................... 82

6.3 COMMUNICATIONS ..................................................................................... 82

6.4 COMMAND AND CONTROL ...................................................................... 86

6.5 SPECIFICATIONS ............................................................................................ 89

6.6 CONCLUSIONS ............................................................................................... 89

APPENDIX 6.1 ............................................................................................................... 91

6.7 BIBLIOGRAPHY ............................................................................................... 93

2
INTRODUCTION

This report presents and describes the Orbiting Plutonian Topographic Image

Craft (OPTIC). The vehicle's twenty year trip will culminate upon arrival at Pluto,

the only major body in the solar system that has not been studied by an earth

launched probe. After arrival OPTIC will begin its data collection which includes

image and radar mapping, surface spectral analysis, and magnetospheric studies.

This initial investigation into the remote study of Plutonian space utilizing an

unmanned probe was conducted by AAE 241 Group 4 at the University of Illinois,

Urbana-Champaign, Spring 1990.

This probe's design was developed based on the Request for Proposal

requirements generated for the spacecraft design section of AAE 241, an

Aeronautical & Astronautical Engineering Senior Design class affiliated with the

Universities Space Research Association. The design work presented herein is the

original work of the six members of Group 4. It has been produced and compiled

based on individual research and knowledge acquired from class work, in addition

to the annotated guidance and information received from outside sources.

Based upon the Request for Proposal emphasis on study of Plutonian space,

and NASA's stress on the importance of not only photographic data, but also

mapping, an orbiter seems to be the best solution. The problems which an orbiter

presents are varied, but all appear solvable. The distinct problems which an orbiter

causes for each subsystem are discussed in their respective sections throughout the

report.

The final design formulation revolved around two important factors: (1) the

ability to collect and return the maximum quantity of information on the Plutonian

system and (2) the weight limitations which the choice of an orbiting craft implied.

The velocity requirements of this type of mission severely limited the weight
available for mission execution - owing to the large portion of overall weight 

required as fuel to fly the craft with present technology. 

While the mission is not constrained to only arrive and examine Pluto, 

Plutonian space is its prime objective. This and other factors, describe within, lead 

to the choice of an orbiting craft. Since the science objectives are what directed this 

mission, the justification for what may appear to be an extravagant task is contained 

within the Science Instrumentation subsection. 

OPTIC DESIGN GROUP 

SCIENTIFIC INSTRUMENTATION ~ r. 

Jonathan E. Kelly /--; C.· -.~,. -r--~ ( 

MISSION MANAGEMENT, PLANNING AND COSTING 

Randall John Hein ~ y-~ 

POWER AND PROPULSION 

David Lee Meyer 

Al I Il'UDE AND ARTICULATION CONTROL / 

David Mark Robinson ;{ ~ ;:ft. 

STRUCTURES 

Mark James Endre 

COMMAND, CONTROL AND COMMUNICATIONS 

Eric W. Summers 

2 
1.0 SCIENCE INSTRUMENTATION

INTRODUCTION

This section presents and describes the four major components of the OPTIC

science subsystem: an electronic imaging unit, a magnetometer assembly, an

ultraviolet spectrometer, and a radar mapping system.

Also included is the explanation of design choice based upon defined mission

objectives.

This subsection was investigated and prepared by Jonathan E. Kelly.

1.1 DIRECTING FACTORS

Pluto is the only remaining planet in the solar system which has not been

studied with the help of an earth-launched spacecraft. What littleinformation that

isknown about thisplanet has been extrapolated from the blurred views which

earthbound telescopesprovide. The quantity and quality of the information thatis

possessed is excellentwhen the limitationsof the collectionprocess and the distance

separating Earth and Pluto are considered.

The successfullaunch and deployment of the Hubble Space Telescope will

increase the knowledge of thistiny world many fold. An actual physical encounter

with Pluto and itssatellite,Charon, utilizinga probe, would increase even that

knowledge by many orders of magnitude. With thisgoal - to investigatethese

bodies intensively- the Orbiting Plutonian Topographic Image Craft (OPTIC) has

been designed.

Since OPTIC's primary objectiveis to gather information about Plutonian

space, the driving factorbehind itsmission design is itsscientificinstrumentation.

3
SCIENCE D_ECTrWES

OPTIC, as the acronym's words imply, is an orbiting probe which will gather

topographic and image data about Pluto. This craft will also collect data concerning

any significant existing magnetic fields that are located around the Pluto system and

additionally perform spectral analysis of the system's bodies' surfaces.

After extensive investigation, which included interviews with a limited

number of the members of the scientific community, and analysis of the

recommendations of past NASA and scientific conferences, an orbiter was chosen as

the best means of obtaining the maximum quantity of useful data and fulfilling the

mission objectives.

During the preliminary investigations for the science subsystem members of

the University of Illinois Department of Geology, all who had previous experience

with planetary studies, were contacted and briefly interviewed. Professor Albert T.

Hsui from this department emphasized that the major questions which a mission to

Pluto should attempt to answer might include: surface makeup, existence of

magnetic fields, gravity, and overall planet size. It was also stressed that

photographic data would most likely help in answering these questions. (4. Hsu/)

In the subsequent study of published documents the importance of images

also became evident. NASA published planetary exploration goals which are in

agreement with Dr. Hsui's recommendations. These primary goals, as listed by the

participants in an Arizona State University (ASU) sponsored Planetary Geology

Workshop, emphasize preliminary acquisition of planet surface characteristics.

Following in the exploration phase should be studies of planets' topography, gravity

magnetic fields, surface chemistry, and mineralogy. This conference emphasized

mapping as the best method for obtaining the maximum information concerning

planetary surfaces. (7. NASA, pgs. 5-7)

4
The present information available about the Plutonian system, coupled with

the round trip signal time, makes a lander unfeasible both technologicaUy and

conceptually. Technologically because the artificial intelligence (AI) necessary for

independent control of a mobile lander at that distance does not presently exist.

That level of AI will probably not be developed suffidenfly by the year 2000.

Even more importantly, with regards to a lander, at the present time,

knowledge of the Pluto system is not extensive enough to provide design criteria for

safely putting a lander (mobile or otherwise) down on the planet. Foremost, there

exists basically zero knowledge concerning surface makeup - a vitally important

piece of information for choosing possible landing sites. Additionally the

knowledge of the physical makeup and conditions of the plutonian environment

does not allow for producing a lander for which survival could be guaranteed with

any confidence.

The flyby option is technologically feasible. It would also be less costly. But

since flyby encounter times are extremely brief (the Neptune encounter of 1989 was

only five days) and one of the main requirements of the proposal is to study

Plutonian space, a probe going solely to Pluto would spend years in transit for an

encounter that would last only days.

Based on the limitations, with regards to a thorough plutonian study, that a

flyby mission has, the logistics and technical problems included in the lander

option, and the recommendations from NASA and the scientific community, an

orbiter was chosen as the best configuration for fulfilling the RFP guidelines.

5
1.2 SCIENCE INSTRUMENTS

CHOSEN SYSTEMS

The most important constraint which an orbiter presents is weight

limitations. Mass is a premium, and, as discussed in the propulsion section, one

kilogram of payload makes necessary an additional 8.4 kilograms of propellant.

Because of this major design limitation, the chosen scientific instruments have been

limited to four. These include: an imaging system consisting of two cameras, an

ultraviolet spectrometer, a magnetometer, and a radar mapper. Table 1.1 lists the

four science instrumentation systems, their weights, and power requirements.

All four of these science instruments will have been flight proven by the

launch of OPTIC. The camera imaging system is modeled after the Cassini/CRAF

imaging systems. The design of the radar mapper is derived from the small radar

mapper used with Pioneer Venus in 1978. Finally the magnetometer and the

spectrometer designs were developed for and placed on Galileo.

TABLE 1.1

SCIENCE INSTRUMENTS

Instrument

Imaging System

Radar Mapper

Ultraviolet Spectrometer

Ma_netometer

Total

Wei_,_t (k_)

36.6

9.7

5.33

5.30

56.93

Power (W)

44.75

18.0

5.33

6.0

74.08

Since mapping and photographic date are of highest priority for NASA in

planetary studies, both the camera system and the radar mapper were first priority

scientific instruments. The camera system is by far the largest and most massive

6
science instrument. It and the other science equipment will be described separately

below.

In the initial studies, a multitude of other possible instruments were

examined. This was limited to five prior to choosing the magnetometer and UV

spectrometer. The eliminated instruments are discussed in the following section.

In the end the choices were made based on science data desired rather than

weight tradeoffs. The weight limitation merely eliminated the number of

instruments, not type.

The magnetometer was chosen for two reasons. The first of those reasons is

the scientific interest in the existence of a magnetic field about Pluto and its satellite.

Information about magnetic fields about a planet was given as much priority by the

ASU conference partidpants as topography and gravity.

The second important reason for the magnetometer links it to the imaging

system in terms of mission success. This linking occurs when the usefulness of the

data collected is considered. Data from magnetometer readings and camera images

is more familiar to more members of the scientific world than other specialized data

acquisition devices. (7. NASA, pg. 13) This means that data collected using these

instruments will be of greater interest to more scientists and, therefore, may spawn

more studies and analysis than other forms of data which may be relevant to only a

few experts.

i

In order to attempt to determine the makeup of the Plutonian surface and its

atmosphere, the UV spectrometer was added as the fourth instrument. The UV

spectrometer fulfills important mission requirements based on the ASU conference

conclusions recommending the investigation of mineralogical and elemental

makeup of planets.

These four instruments are at present the only proposed science systems for

OPTIC. As emphasized previously, mass limitations called for a compromise which

7
would result in minimum weight with maximum useful scientific data collection.

It is felt that these instruments wiU provide a wealth of information about Pluto.

Their capabilities should allow OPTIC to fulfill the NASA./scientific request for

visual images, topographical data, magnetic field information, and surface

chemistry/mineralogy of Pluto and to a lesser degree its satellite Charon.

OTHER INSTRUMENTS CONSIDERED

Following are the major science instruments which were cut from the OPTIC

science package: a laser altimeter, a gravity analyzing microprobe, and an infrared

spectrometer. These three instruments were ruled out primarily because of the

uncertainty of either the need for them, or the ability to develop them. Thus, they

were eliminated due to these constraints, combined with the concern for mass,.

The most promising and valuable of these is probably the gravity analyzing

microprobe. This concept envisions a microprobe mounted transponder, ejected

from the main probe, and tracked using OPTIC's secondary antenna. Data collected

concerning the microprobe's motion could be used for gravity field calculations for

the Pluto - Charon system. Since this is an untested/developed concept it was

shelved to conserve weight and save development costs. If significant advances in

propulsion, trajectory, or budget, are made during the following development stage,

more investigation in this instrument is recommended.

The laser altimeter is an exciting topographical data collection concept

planned for the Mars Observer. (6. Mars Observer, pg. 79) Unfortunately, this

instrument requires an extensive framework of information concerning the target

body's gravitational field to function. The necessary data concerning Pluto is not

available. Without this data, preflight calibration of an altimeter of this type would

be nearly impossible.

8
Finally,the infrared spectrometer is an instrument which has travelled on

numerous interplanetarymissions before. Itsmain task has been the study of

appreciable atmospheres. Since itis not believed that Pluto has much atmosphere

in existence,the trade off for saving mass seemed a betterproposition.

1.3 INSTRUMENT DESCRIPTIONS

IMAGING SYSTEM

The imaging system for OPTIC utilizesthe Imaging Science Subsystem (ISS),

originallydeveloped for the Comet Rendezvous Asteroid Flyby (CRAF) and Cassini

missions. This two camera system provides OPTIC with a reliable,flighttested

(assuming both missions are executed prior to 2004), system that will not incur the

extravagant costs which new system design implies.

The ISS has been designed for use on several missions. Ithas been developed

in such a manner that itcan easilybe adapted to thismission to Pluto. The systems

design provides for differentdata output rates,distinctdata compaction options, and

the abilityto be used for navigation purposes.

ISS employs a wide angle camera and a narrow angle camera. Both of these

draw upon a common electronicsmodule. The relevant data for both cameras is

outlined in Table 1.2. The two focallengths provide for two. distinctscalesof image

resolution. Each contains filterswhich allow for varied spectralstudy of their

focused target.

9
Camera

Optics Type

FocalLensth

SpectralRange

Filters

Field of View

TABLE 1.2

CAMERA DATA

(Adapted from 5. J'PL, pgs. 8,

Narrow Angle

Ritchey Chretien

2000 mm

200-1100 nm

22

0.35 ° square

9)

Wide Angle

Refractor

250 mm

350-1100 nm

14

2.8 ° square

The imaging system cameras can operated simultaneously and they can be

calibrated and ready for use in less than an hour. The system is equipped with

automatic exposure control and the frame time rates can be varied from 9 to 1479

seconds. The focusing ability for the Cassini/CRAF versions allow for 25 and 3.8 km

passes, respectively. (5. JPL, pg. 8) While present planning for OPTIC place these

altitudes well below the mission orbits (~ 2 Pluto radii from the surface) the 25 km

value allows for great flexibility in imaging mission modification. (The 3.8 km

value for CRAF is achieved with extra lenses). (5. JPL, pg. 8)

The frame time ranges are the total system process time. This includes

exposure time and the time needed by the system to prepare for the next

exposure. (filter rotation, etc.) The automatic exposure control is accomplished by

taking two photographs. The first is used to supply irradiance information for the

control system. The system then recalibrates and takes the second photo, which is

the science image data. (5. J'PL, pg.10)

The imaging subsystem weighs 36.6 kilograms and consumes a maximum of

44.75 watts of power. (1. Advanced Projects) They are both mounted on the three

degree of freedom science platform. This provides for nearly unconstrained aiming

possibilities. In this mannerboth cameras can be used to photograph Charon and

any other targets of opportunity. The cameras can be activated in route for asteroid

I0
study, Jupiter analysis, and navigation backup. Figures 1.1 and 1.2 show the side

view representations of both cameras, including their dimensions.

FIGURE 1.1

NARROW ANGLE CAMERA

(Adapted from 5. J'PL fig 3-2)

I

v

40 (:n't

! I I

FIGURE 1.2

WIDE ANGLE CAMERA

(Adapted from 5. yPL fig 3-3)

I I

ID(I I

_CrN

_cm

11
IMAGING PROCESS

While the mass of Pluto is known, there remains some uncertainty

concerning the value radius. Most recent information indicates a radius of 1150 krn.

(8. SAIC) The program used for thismission design trajectorycalculations,

(seesection 2),assumes a radius of 1500 km. This discrepancy presents

problems for developing exact mapping times.

While the finalmapping orbitisstated as an ellipsewith a periapse of 2

Plutonian radiifrom the surface,thisisa 700 km differencein altitudedepending

on the assumed radius. Since alltrajectorymechanics are based on the 1500 km

radius, thisvalue was used for preliminary mapping calculations.

Using the narrow angle camera, and assuming 100% coverage with minimal

overlap, photographs at orbitperiapse would cover 336 square kilometers (see

Appendix 1.1). Mapping is assumed to occur while the true anomaly ranges from

270 ° to 90 °. this provides a pass time of 5.4973 hours. With this period of mapping

time per orbit the total mapping duration is calculated to be 482.2 days. 17.46%, Or

84.2 days, of this time is actually spent imaging the planet. These numbers are

obtained using a photographic rate of 68.6 seconds per photo.

The ISS provides variable data rates to which the photo rate can be fitted.

Using the available compression rate of 2 to 1, and assuming the automatic exposure

control is activated (requiring two photos for every one science image), the frame

time is 60 seconds. This mapping rate generates 175 kilobits per minutes of output

while providing the highest resolution.(5. JPL, pg. 12) Exact resolution in units of

size depends on altitude, but the narrow angle camera has a resolution of 6

microradians square and the wide angle 48. (5. JPL, pg. 8)

If upon arrival, fuel stores are sufficient orbit altitude can be decreased, and a

slower rate (frame time) could be utilized. The slower output rate means lower

12
resolution, but combined with the lower altitude, can provide near equal image

resolution. The flexibility of the SSI allows OPTIC to proceed towards Pluto even

without full knowledge of final science mission conditions. If, after arriving at

Pluto, initial data collection can be used by scientists and controllers to adjust the

final imaging course to an optimal route.

MAGNETOMETER

OPTICs magnetometer is based upon the Galileo magnetometer design.

When magnetometer data is to be collected the craft is put into a rotation rate of 3.15

rpm around its Z-axis. This spin is induced to obtain more totally encompassing

field data on any existing magnetosphere.(9. Yeates, pg. 105) Because of this

spinning motion, magnetometer data acquisition cannot occur simultaneously with

the other instruments' data coUection. This instrument is to be used in an attempt

to answer questions pertaining to the existence of a magnetosphere about the planet.

The instrument's sensors, (see Figure 1.3) separated into two sets of three, are

located at the end of the 10 meter boom. Separated slightly to provide correction for

any magnetic fields generated by OPTICs other electronics, the two sets measure

magnetic fields of distinct intensities. The farthest set is sensitive over the span of

_+.32 to .+.512 nT, and the second set from :b.512 to +16384 nT.(9. Yeates, pg. 131)

13
FIGURE 1.3

MAGNETOMETER SENSORS

(Adapted from 9. Yeates, pg. fig. 93 a.)

As the sensors move through space, both forward along the orbit trajectory,

and about with respect to OPTICs Z-axis, an analog voltage is generated

propor6onal to the magnetic field. This is converted to a 16 bit digital signal by the

magnetometers data system. The data system samples, averages, and stores the

measured data prior to its transfer to the main OPTIC computer. (9. Yeates, pg. 133)

From the main computer the data is transmitted to earth.

The data can either be processed and stored in "packets" that are partitioned

by equal increments of time over the duration of the measurement period, or in a

form of X,Y, and Z location coordinates separated data. There are also two lesser

used processing modes involving the extremes of long duration measurements and

short high speed data acquisition. When not operating a final mode merely places

the system on hold for command changes. (9. Yeates, pg. 133)

The magnetometer weighs 5.3 kg and utilizes approximately 6.0 W of

power.(3. Giampeoli) Data acquisition will be performed during each distinct orbit

(arrival and mapping) in order to analyze the broadest reaches of the Plutonian

magnetosphere. To conserve fuel in the control thrusters and increase coverage, the

14
instrument will be activated during the lastorbitsof the arrivalorbitpath, remain

on during the orbit change, and run during the initialpasses on the mapping orbit.

The system can also be activated to collectmore Jovian magnetic data as OPTIC

nears the planet for itsgravity assistduring the tripout.

ULTRAVIOLET SPECTROMETER

This instrument was also originallydesigned for the Galileo probe. For

Galileo'smission itwas used to analyze the Jovian atmosphere and itssatellites'

surface makeup. Without the highly significantatmosphere of Jupiter'ssystem to

analyze, OPTICs UV spectrometer will be utilized to perform detailed analysis of

the Plutonian surface. It will, additionally, if arrival orbit orientation permits, focus

upon Charon's surface.

The spectrometer extends OPTICs spectral range from the 1100 angstroms

covered by the ISS to include the span between 1150 angstroms to 4300 angstroms.

Its observations will provide ultraviolet reflectivity readings. This data is pertinent

to the physical state in which the materials on the surface exist:, i.e. ice, frost and

grain size. It will also attempt to detect the presence of the elements hydrogen,

oxygen, and nitrogen, in their atomic states.(9. Yeates, pg. 130)

The system has four major components which are housed within the

instrument: a telescope, a monochromator, three detectors, and the system control

logic unit. The components are identified in Figure 1.4.

15
FIGURE 1.4

SPECTROMETER ASSEMBLY

(Adapted from 9. Yeates,fig.91)

24em

Ik _ E CTORHEADS

CONTROL LOGIC

10em

The telescope has a 250 mm aperture which creates a field of view of 0.1 ° by

1.4 ° for the 1100 to 1900 and 2.800 to 4,300 angstrom detectors. The field of view for

the 1600 to 3000 angstrom range is 0.1 ° by 0.4 °. These field of views provide for

spectral analysis of small selected regions upon the Plutonian surface.(9. Yeates, pg.

130)

The monochromator, with a reflecting diffraction grating of 125 mm focal

length, disperses the ultraviolet light. Grating position (and, therefore, wavelength

measured), is regulated by the control logic of the system which instructs the grating

drive. The resolution is 13 angstroms in the f'trst order spectrum, and 7 angstroms

in the second order spectrum.(9. Yeates, pg. 131)

Photons which hit the 3 detectors produce pulses which are counted and read

every 7E-4 seconds. In turn, these pulses are sent to the main computer for

transmission back to earth.(9. Yeates, pg. 131)
The system processor can instruct the detectors to measure for only one given

wavelength or view the entire UV spectrum approximately every 4 1/3 seconds.

These, and the variations available in between allow investigations that range from

single wavelength intensity changes across a large planetary swath, or broader

general analysis.(9. Yeates, pg. 131)

The system weighs 5.33 kg and is run with 5.33 W of power.(9. Yeates, pg. 131)

The spectrometer is mounted on the scan platform and runs simultaneously with

the ISS providing a wide range of spectral coverage for all imaged targets.

RADAR MAPPER

The radar mapper of the OPTIC science subsystem is the most modified

instrument to be utilized.This small mapper uses the secondary antenna to obtain

topographic data of the planets surface. The radio science possible with thismapper

includes the search for rings (allthe outer planets have been found to have rings,

with the exception of Pluto) and precise radiimeasurements. These are

accomplished using occulation measurements of the received signals on earth. The

extreme distance to Pluto make the success of thistype of testimprobable.(9. Yeates,

pg. 55)

In upgrading the Radar mapper, the basic design from the Pioneer Venus craft

will be modified to transmit on two bands rather than one. The addition of the 3 cm

X-band, to supplement the 12 cm S-band, will provide better,more complete radar

images of what is assumed to be a mostly rock surface.

The second main modification involves the use of the 1 m diameter

secondary antenna dish. The Pioneer Venus version used a 0.38 m diameter dish.

This increase in size will increase the overall topographic resolution of the data

collectedby the mapper.

17
The final difference in the OPTIC version is in its use. This model will be

operated continuously during the photo mapping passes, rather than on a rotating

basis as on the Pioneer Venus.

Utilizing the collected radar data, in conjunction with the data that will be

obtained concerning Pluto's exact dimensions, it is hoped that absolute surface

elevations can be calculated.

The Pioneer Venus system weighed 9.7 kg and used 18 W of electrical

power.(2. Fimmel, pg. 58) It is assumed that the present technology may lower

these values, even with the upgrading being planned,but the Pioneer Venus

numbers have been used for all OPTIC system calculations.

18
APPENDIX 1.1

MAPPING CALCULATIONS

Asstmunptions from MUIaMP (8. SAIC):

rpl = 1500 km P" pl = 663.5622 km^3/s ^2

Calculations and Data:

ApI = 2.8274E7 kin ^2

mapping

orbit: rperiapse= 3rpl

rapoapse = 5rpl

Tmap orbit = 2x(aA3/_t pl)A1/2

= 113361.589 s

= 31.4893 hr

Partialorbit period given by: t=(aA3/p, pl)^l/2[E-esinE] where E, eccentric anomaly, is

defined as: E=2atan[((1-e)/(l+e))A1/2tan(f/2)] f = true anomaly

Mapping occurs from f = 270 ° and f = 90 °

These equations yield a mapping pass time of 5.4973 hours.

19
APPENDIX 1.1 (con0

Figure 1.5

MAPPING GEOMETRY

eld of view:

=0135°

altitude

X

x = ztan_

=9.163 km

Narrow Angle Camera field of view: 0.35 ° square 0.35/2=t3 (see figure 1.5)

Minimum coverage occurs at periapse (3000 km). One mapping pass covers one half

of Pluto's circumference (4712.3890 km)

This yields a total area of 335.8428 km ^2. This value is the minimum

coverage per photo. Dividing ApI by this value yields the number of photos

necessary for complete coverage: 84188 photos. Using the distance covered in one

pass, 4712.389 kin, and the coverage of each photo the value of 257 photos per

mapping pass this reduces to one photo every 77 seconds. Actual mapping time

becomes 75 days of actual photographing, or a total time to map of 430 days

fincluding orbit time from f = 90 ° to f = 270 °)

20
1.5 BIBLIOGRAPHY

1. Advanced Projects Division, Jet Propulsion Laboratory, California Institute of

Technology, Pasadena, California.

2. Fimmel, Richard O., et. al. Pioneer Venu3, U.S. Government Printing Office:

Washington D.C. (1983).

3. Giampeoli, R. "Galileao Spacecraft Mass Report and Equipment List", Issue #36,

May 1988.

4. Hsui, Albert T. Associate Professor, Department of Geology, University of

IUinois. Interview conducted February 23, 1990.

. J'PL, Announ¢¢m¢nt of Opportunity. OSSA-I-89, Cassini Mission, Jet Propulsion

Laboratory, California Institute of Technology, Pasadena, California. October

10, 1989.

. "Mars Observer Begins New Era Using Proven Spacecraft Design." Aviation

Week and Space Technolok_¢, Vol. 131, pg. 79, October 9, 1989.

. NASA Office of Space Science and Application. Planetary_ Geology: Goals,

Futur_ Direction, and Recommendations, Report from the Planetary Geology

Workshop, Arizona State University, January 1987, U.S. Government

Printing Office: Washington D.C. (1988).

8. Scientific Applications International Corporation (SAIC), MULMP program

data. 1990 version.

9. Yeates, C. M., et. al. Galileo: Exploration of Tupiter's System. U.S. Government

Printing Office: Washington D.C. (1985).

21
2.0 MISSION MANAGEMENT, PLANNING AND

COSTING

INTRODUCTION

The primary requirement and problem for an orbiter is the need to keep the velocity

at Pluto a minimum. If the velocity is too high, the mount of fuel needed for orbit

insertion becomes unrealistic. Also of concern is the need to keep the delta-v at Earth

departure at a minimum while reducing mission time as much as possible.

This section was prepared and written by Randall John Hein.

2.1 METHOD OF ATTACK

To find the optimum orbits a technique using synodic periods was used. Each pair

of planets in a trajectory (exduding Pluto) were examined. The optimum transfer point

was propagated through our 10 year launch window. When only 2 planets (excluding

Pluto), such as Earth and Jupiter, are involved in a trajectory, a mission time frame of 10 to

24 years was examined in each synodic period. We see from Figure 2.1 that the optimal

launch date will be about December 14, 2004 or November 13, 2003, depending on mission

length. When 3 or more planets are involved, such an in a Earth-Mars-Jupiter-Pluto

trajectory, the optimum points in each pair of synodic periods were compared until one

ore more viable launch dates could be found. Each of these was then examined for a 10 to

24 year mission.

22
3O

>,

FIGURE 2.1

FLIGHT TIME VS. TOTAL DELTA-V

FOR JUPITER PLUTO TRAJ'ECTORY

WITH ORBIT INSERTION AT PLUTO

LAUNCH AUG 9, 2000

LAUNCH SEP 10, 2001

'LAUNCH OCT 12, 2002

LAUNCH NOV 13, 2003

LAUNCH DEC 14, 2004

JAN 15, 2006

LAUNCH FEB 16, 2007

LAUNCH MAR 19, 2008

APR 20, 2009

10 , •

5 15 25

FLIGHT TIME (YEARS)

Many possible mission plans were examined.

the Galileo mission(1. D'Amario) were tried first.

Trajectories similar to that used in

The Venus-Earth-Jupiter-Pluto

trajectory proved to yield a better initial delta-v than the a Jupiter-Pluto trajectory (fig 2.2),

but the final delta-v was too high (fig. 2.3). The same problems accrued with other

trajectories that use inner planets for a gravity assist as opposed to strictly using outer

planets. A Mars-Jupiter-Pluto trajectory was also considered. Though the total and initial

delta-v's were the best (fig. 2.4 and 2.2), the Jupiter-Pluto trajectory gave the better final

delta-v (fig 2.3).

23
FIGURE 2.2

COMPARISON OF INITIAL DELTA-V'S

FOR PLUTO MISSION WITH ORBIT INSERTION

9

8

_7

;6

4

3

0 10 20 30

JUPITER-PLUTO

VENUS-EARTH-JUPITER-P

MARS-JUPITER-PLUTO

FLIGHT TIME (YEARS)

FIGURE 2.3

COMPARISON OF FINAL DELTA V's

FOR PLUTO MISSIONS WITH ORBIT INSERTION

3O

>,

= lo

0 | I

15 25

,.--..ffi..- JUPITER-PLUTO

VENUS-EARTH-JUPIT[

MARS-JUPITER-PLUTC

FLIGHT TIME (YEARS)

24
FIGURE 2.4

COMPARISON OF TOTAL DELTA V's

FOR PLUTO MISSIONS

WITH ORBIT INSERTION

3O

0

5 15 25

JUPITER-PLUTO

VENUS-EARTH-JUPIT_

MARS-JUPITER-PLUTC

FLIGHT TIME (YEARS)

Saturn alone was not a useful gravity assistbody. The only possible advantage to

using Saturn would be in a Saturn-Jupitertrajectory.Unfortunately, Saturn lags Jupiter's

orbit during our prescribed launch window. A similarproblem occurs when using

Neptune or Uranus for a gravity assisteddeceleration.Pluto'sorbit lags Neptune's and

Uranus', thus eliminating the option.

The Jupiter-Plutotrajectorywas finallychosen because of the need for a low final

delta-v. The decision for a 20 year mission came from the need to keep the fuel weight for

the finaldelta-vunder 3000 kg. By looking at figure 2.5one can see the that a mission time

of 20 years or more isrequired to keep the finaldelta-vunder the needed 5 km/s.

25
FIGURE 2.5

FLIGHT TIME VS. FINAL DELTA-V

COMPARED WITH

FUEL MASS VS FINAL DELTA-V

30

28

i 26

24

22

p, 20

:=

16

14

12

-i/\r

"N

2 4 6 8 10 12

FINAL DELTA-V (KM/S)

6000

4000 ...

v

.¢

20O0

,0

14

FLIGHT TIME (YEARS)

MASS (KG)

DELTA-V REOUIREMENTS

A Jupiter-Pluto trajectory was chosen with a gravity assist at Jupiter. Once in Low

Earth Orbit (LEO), 300 KM above the Earth's surface a delta-v of 7.116 krn/s would be

needed for orbit departure. At 576.5 days into the mission OPTIC would reach its close

approach point of Jupiter at 2,196,911 km from the planet's surface. At the 20 year mark, an

orbit insertion burn (delta-v - 4.927) will be needed. The orbit radius at perhaps will be

4500 krn and 15000 km at periapse. Upon completion of required data acquisition

(approximately 17 days) a second burn will take place (delta-v = .0470 km/s) at periapse.

This will put OPTIC in an orbit with a radius of 7500 km at apoapse and 4500 km at

26
periapse. The reason for the non circular orbits is to reduce the delta-v needed for orbit

insertion.

DECEMBER 13, 2004

Launch OPTIC aboard Shuttle-C, using an initial Space Transfer Vehicle (ISTV)

for departure from LEO.

JULY 12, 2006

Jupiter Encounter: OPTIC will study Jupiter for 5 days while using the planet

Gravity assist.

for a

December 13, 2024

Pluto Encounter: Optic will do an orbit insertion burn to get into a eccentric orbit

about Pluto. The time in this orbit will be used for scientific study focussed at Charon.

JANUARY 1,2025 [ This date isdependent on scientificneeds.]

Orbit Maneuver: OPTIC will move into a lower orbitabout Pluto to allow for

mapping and furtherscientificstudy of Pluto.

DECEMBER 13, 2026 [This date is dependent on scientific needs.]

Mission ends.

2.2 COSTING

The cost of OPTIC comes to $999.78 Million in fiscal 1989 dollars. An additional

charge for the ISTV booster (cost unavailable) and Shutfle-C must be included. The

estimated cost for the Shuttle-C is about $2000 per pound which comes to about $69.4

Minion(2. Kolcum, p.134). To keep the cost down, we used as many unmodified

components as possible. Due to the high delta-v required for orbit insertion, a propulsion

system requiring major modifications had to be developed. Several other subsystems were

forced to do major modifications or design new components due to the duration of the

mission and the distance from Earth.

27
APPENDIX 2.1

FORMULA_ USED FOR COSTING(costs are given in fiscal 1977 dollars)

DEVELOPMENT PROJECT - FLIGHT HARDWARE

STRUCTURES & DEVICES

DLH = 1.626 (N'M)^0.9046

Thermal Control, Cabling & Pyrotechnics

DLH = EXP(4.2702+.00608*N*M)

Propulsion

DLH = 56.1878(N'M)^0.4166

Attitude & Articulation Control

DLH = 21.328(N'M)^0.72,30

Telecommunications

DLH = 4.471(n'm)^1.1360

Antennas

DLH = 6.093(N'M)^1.1348

Command & Data Handling

DLH - exp(4.2605+0.02414*N*M)

RTG Power

DLH = 65.3(N'M)^0.3554

Power

DLH = EXP(3.9633+0.00911*N*M)

Radar

DLD = 11.409(N'M)^0.9579

Imaging

DLH = 4.463(N-M)^1.0369

Particle & Field Instruments

Labor hrs. to labor cost = 34.52

RLH = 1.399(N'M)^0.7445

Labor hrs. to labor cost = 34.03

RLH - 3.731(N'M)^0.6082

Labor hrs. to labor cost = 38.11

RLH ---(N'M)^0.9011

Labor hrs. to labor cost = 35.58

RLH = 1.932(N'M)

Labor hrs. to labor cost = 33.49

RLH = 1.626(N'M)^1.1885

Labor hrs. to labor cost - 34.52

RLH -- 3.339(N'M)

Labor hrs. to labor cost = 30.61

RLH = EXP(2.8679+.02726*N*M)

Labor hrs. to labor cost = 30.21

RLH = 7.88(N'M)^0.7150

Labor hrs. to labor cost = 32.77

RLH --- EXP(2.5183+.01204*N*M)

Labor hrs. to labor cost = 31.83

RLH - 1.2227(N'M)^1.2367

Labor hrs. to labor cost = 34.14

RLH = (N'M)^1.1520

Labor hrs. to labor cost = 36.05

28
DLH = 25.948(N'M)^.7215

Remote Sensing Instruments

DLH -- 25.948(N'M)^.5990

RLH - 0.790(N'M)^1.3976

Labor hrs. to labor cost ---35.0

RLH - .790(N'M)0.8393

DEVELOPMENT PROJECT - SUPPORT FUNCTION

System Support & Ground Equipment

DLH = .36172(SUM DLD Hardware) Labor hrs. to labor cost =32.45

Launch +30 days Operations & Ground Software

DLH - 0.09808(SUN DLH Hardware) Labor hrs. to labor cost ---34.42

Imaging Data Development

DLH -- .00124(Pixies per line)^1.629

Science Data development

Labor hrs. to labor cost = 35.87

Labor hrs. to labor cost = 50.87

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

Program Management Labor hrs. to labor cost = 31.07

DLH -- 0.10097(SUM DLH all categories)^0.9670

FLIGHT PROJECT

Flight Operations Labor hrs. to labor cost = 33.90

DLH ---((SUM DLH Hardware/3100)A.6)*(10.7MD + ED)

Data Analysis Labor hrs. to labor cost - 35.76

DLH = 0.425(DLH Flight operations

COST REDUCTION ALGORITHM BY INHERITANCE CLASS

X1 = % of subsystem off- the-shelf

X2 = % of subsystem exact repeat

X3 = % of subsystem minor modifications

X4 = % of subsystem major modifications

X5 = % of subsystem new design

Z -- 1(X1) + .8(X2) + .25(X3) +.05(X4) + 0.0(X5)

Total costs = (100%-z)NRC + RC

' 29
NRC = Non-recurring cost = (DLH-RLH) * Labor hrs. to total cost

RC = Recurring costs = RLH * Labor hrs. to total cost

CONVERSION FROM 1977 DOLLARS TO 1989 DOLLARS

Cost in 1989 - Cost in 1977 * (894.7/505)

TRAJECTORY PLANNING

All figures were derived using MULIMP (2. Frielander)

30
2.3 BIBLIOGRAPHY

1 D'Amario,Louis A "Galileo 1989 VEEGA Mi_i0n D_scription" AAS/AIAA

Astrodynamics Specialist Conference, August 7-10, 1989

2. Frielander, Alan L. "MULTI-IMPULSE TRAFECTORY AND MASS

OPTIMIZATION PROGRAM", Science Applications, Inc 1974

e Kolcum, Edward H. "Decision on Shuttle-C Expected After Internal NASA

Review," Aviation Week and Space Technology. June 19, 1989, pp 123-125.

31
3.0 POWER AND PROPULSION

INTRODUCTION

The following section will explain the design process of both the power and

propulsion subsystems of OPTIC. This section was made especially difficult because

OPTICs mission plan calls for art orbit insertion about Pluto after a 20 year voyage.

These two criteria put an incredible demand on both subsystems. This section was

prepared by David L. Meyer during the spring of 1990.

3.1 POWER SUBSYSTEM

The furst problem that needed to be addressed was the selection of the power

system. This process was simplified because of the type of mission. Solar ceils were

ruled out immediately because of the distance from the sun that OPTIC will be

travelling. The power that solar ceils can produce greatly diminishes past a distance

of 2 A.U.'s (-3E+8 kin) from the sun. The fact that OPTIC will be orbiting Pluto at a

distance of approximately 40 A.U.'s makes the use of solar ceils impossible because

the power at just 6 A.U.'s is reduced to about 5-10% of the power available at 1 A.U.

(5. Koepke, p 11) With solar cells an impossibility, the only alternatives are nuclear

power systems.

Incorporating a nuclear power system into OPTIC will most likely bring out

some political opposition. The groups that presently oppose the use of nuclear

power in space (SANE/Freeze, Citizens to Stop Plutonium in Space, and the

National Mobilization for Survival) will probably still be active well into the

twenty-first century. Their argument is based around two possible disasters: an

explosion at launch and a possible reentry during a fly-by of Earth. The possibility of

32
an explosion during launch has thoroughly tested for in existing Radioisotope

Thermoelectric Generators (RTGs). The RTGs have been subjected to both loads of

up to 2000 psi and projectiles traveling at

TABLE 3.1

PREDICTED RTG ADVANCEMENTS

.qollrc_ _r_pcqficnowpr ('W/k_ Dafp

t Koepke 11 1990

Schock --10.5 "future"

Mondt 10-13 2000

speeds of 360 m/s. The Challenger accident resulted in loads of only 10 psi and in

the case of a solid rocket booster exploding, the shrapnel would be traveling at

approximately 90 m/s (9. Nichols, pp 8-15). This data suggests that RTGs should be

100% safe during a launch. This leaves the case of a possible reentry into Earth's

atmosphere during a fly-by. Our trajectory calls for OPTIC to leave low-Earth orbit

(LEO) and to never return near Earth-space. Given existing test results and the

trajectorythatOPTIC will take,nuclear power willbe a very safe option.

Now that the power source wili be nuclear power, the issue iswhether to use

a nuclear reactor or RTGs. The problem encountered in the past with nuclear

reactorshas been theirextremely high mass and need for shielding. Recently, the

mass of these reactorshas been rapidly decreasing with technologicaladvancements.

Ithas been forecasted that the specificpower of nuclear reactorscould reach as high

as 55 W/kg by 1991 (5. Koepke, p 17). However, at thisdate,no nuclear reactors

have been flighttested.

Presently,RTGs have a relativelylow specificpower of about 5 W/kg (5.

Koepke, p 17). According to a number of sources,specificpower of RTGs could reach

as high as 13 W/kg (seeTable 3.1).

Although nuclear reactorswiU have superior specificpower, other factors

must also be weighed. Our mission duration isover 21 years. This figure does not

33
include any time in storage or grounded because of various reasons. The time at

which the demand on the power supply will be the greatest is when it will be

needed the most. This time will be over 20 years after the launch date while OPTIC

is in orbit about Pluto. The mission will be a complete waste if the power system is

not working properly after 20 years. Knowing this, design lifetime, past results and

reliability play a much larger part in the power system design. For this mission,

those criteria outweigh specific power.

By the year 2000, design lifetime for nuclear reactor thermoelectric conversion

will be from 10--14 years with 10 years being the most "probable" value (8. Mondt, p

40). This is less than half of our mission duration and there is no flight-tested data

to suggest a longer lifetime.

The current design lifetime for RTGs is about 5 years. Lifetime predictions for

the year 1990 are roughly the same as for nuclear power generators (-10 years). This

figure will most likely keep demonstrating an upward trend towards 1999 (8.

Mondt, p 47). However, what separates the 10 year design life of RTGs from the 10

year design life of nuclear reactors is the flight-proven data from various missions

where RTGs were used. RTGs that have been used in the past have provided both

longer lifetimes and power that was weft above the predicted values (1. Bennett, p

327). Some good examples of this are the Voyager probes that were launched in the

early 1970's and are still operating. Conditions that RTGs must undergo can also

affect the lifetime. For example, reducing the hot junction operating temperature of

the RTG assembly can prolong the life of the RTGs substantially (8. Mondt, p 48).

Since solar cells are not feasible, the decision of which power system was

decided by the trade of mass for dependability and longer life. The fact that an

orbiter is being attempted requires that weight be minimized as much as possible.

However, a light power system that does not work upon arrival cannot compare to a

heavier power system that is dependable.

34
FIGURE 3.1

RELATIONSHIP BETWEEN

PROPELLANT MASS AND RTG MASS

2600

2.50o

24oo

z oo

21oo

2oo0

1900

• °

I I I

0 2o 40 60 80

mass of RIGs (kg)

Now that a power source has been chosen, ithas to be designed around the

power draws of the other subsystems. The specificpower of the RTGs, for design

purposes, was setat 12 W/kg. As statedearlier,predicted values for specificpower

by the end of the 20th century are 10-13 W/kg. Although 12 W/kg isnot the most

conservative estimate, itisjustabove the middle of the predicted range, so itshould

not be very faroff (ifany) from the actualvalues. As a resultof thishigh specific

power, no batterieswere required because the various subsystems will be able to

draw power offof the main power source without an appreciable weight loss. The

lack of batteriestakes a lotof the complexity out of the power system. With a

redundant power distributionfrom the RTGs, thisshould be a very simple and

reliablepower system while, at the same time, not being overly massive.

The members of the design team were told to keep power draws down to a

minimum. Figure 3.1 shows the relationshipof the mass of the RTGs to the mass of

35
propellant needed to achieve the orbitinsertioninto Plutonian space.

relationshipislinear,itishardly a 1:1ratio. Every

Although the

TABLE 3.2

t Scientific instrumentation i

Articulation & control

ommand & communications

Structural

SUBSYSTEM POWER DRAWS AND MASSES

_1_h._tv.qfern rvfax. Powpr (I/V_

61.38

20.0

10.0

145.0

Mass (k¢)

v

56.93

68.1 (inculdes 40 kg of Hydrazine)

67.0

80.0

kilogram of RTGs added to the power system adds 7.73 kg of propellant which,

inturn, adds .67 kg of tank weight. Therefore, for every kg added to

the power system, 8.4 kg of extra mass is added to OPTIC. Relating these figures to

Watts, every Watt of power needed for a subsystem adds .78 kg of mass to OPTIC.

The above information illustrates how important rationing of power is. The final

values of subsystem power draws, along with masses, are shown in table 3.2.

The degradation rate of the RTGs had to be determined in order for OPTIC to

be operational at Pluto. A graph relating the ratio of power output to original power

to time was available in Bennett et al. on page 327. This contained the following

data: LES pre-launch prediction, LES actual data, Voyager pre-launch data and actual

Voyager data. The actual data of both cases was much better than the predicted

values. There was also a substantial improvement from the LES results to

Voyager's results. The plot covered a time period of 44000 hours. The curves were

very conservatively extended to 240000 hours (27.38 years). Included in these 27.38

years are the 20 year mission time and an allowance of 5 years in storage or

grounded preceding the launch. This leaves 2.38 years at Pluto. Information from

the Scientific Instrumentation subsystem states that it will take 1.18 years to

completely map Pluto. Excluding the pre-launch allowance of 5 years, this is a

36
cushion of about I year. After 240000 hours, the ratio of power output to original

power is at .770 for actual Voyager data and .722 for predicted Voyager data. For

design purposes, the ratio of .77 will be used. This is highly conservative because

the curve was extrapolated conservatively and it is based on out-dated technology.

However, because of the long mission duration, conservatism is best.

Using the total peak power of 236.38 W, dividing by 12 W/kg and then

dividing by .77, the mass of the RTGs that need to be installed initially into OPTIC is

25.582 kg. This is a starting power of about 307 W. The power that will be left on

OPTIC, once in Plutonian space, will be able to keep every instrument running

simultaneously for 2.38 years. This substantial time cushion, along with the

conservative estimate of the available power, should keep OPTIC operational for

many years after orbit insertion.

Additional components for the power system are as follows: a power

regulator, a power control unit and a power distribution unit. It was assumed that

each subsystem would provide their own inverters. The three components

mentioned above were taken straight from the equipment list of the current Galileo

probe. The reasoning behind this was that these devices had to be of the highest

quality and dependability because of the myriad of

electronics and experiments on board Galileo. Although OPTICs equipment list is

minimal, dependable, heavy-duty electronic regulators and distributors are needed

to maintain a steady power signal throughout the full duration of the mission.

Another positive factor of the Galileo equipment is that there will be some flight

data which could point out potential problem areas before installation into OPTIC.

The need to minimize weight was explained above. Since the electronic

subsystem is the heart of any mission, any failures could severely jeopardize the

mission's success. The one problem area that the power system might encounter

are breakdowns because of the long mission duration. So, when selecting

37
components and a power supply, reliability was stressed more than weight savings.

Even after these precautions to guard against any failures, a redundant system was

attempted (as shown in figure 3.2). In sum, barring any catastrophic failures, the

power subsystem should provide reliable power and it will most likely outlast

OPTICs mission time.

FIGURE 3.2

SCHEMATIC OF POWER SYSTEM

,m

I"--"-- Power Power

'_ Gs Regulator Power Control Dist. Unit Unit

E3 E3 E3 E3 @@@

various instruments

3.2 PROPULSION SUBSYSTEM

Similarly to the power subsystem, the first problem encountered with the

propulsion subsystem was the selection of a propellant. Because our mission calls

for an orbit insertion about Pluto, there will be a great demand on the propulsion

system. The change in velocity (Av) needed at Pluto is 4927 m/s. The mission

duration is about 20 years. This combination of a high Av and the need for longterm storage presents quite a problem.

If the high Av was the main focus, a propellant with a high specific

thrust (Isp) would be the best choice. Similarly, if the mission duration was

the only aspect to be analyzed, a highly storable propellant would be chosen.

38
The combination of these two parameters in an existing propellant is hard to find.

Most stable propellants have relatively low Isp values and the high performance

fuels are not storable for long periods of time (12. Sutton, pp 168-182).

FIGURE 3.3

PROPELLANT MASS NEEDED FOR

Av OF 4927 m/s vs. Isp

12OOO

8OOO

point 60OO

0

240 260 280 300 320 340 360

specific thrust (s)

More preference was given to stability because if the propellant has boiled off

or eaten away the tanks before arrival at Pluto, there would be no orbiter. However,

the storable fuel must have a good performance value so the tank size does not

become infeasible.

The question of whether to use a monopropellant or a bipropellant also

needed to be addressed. Monopropellants have the advantage of simplicity while

giving up some performance values, as compared to bipropellants. One of the most

common monopropellants, hydrazine, has a theoretical Isp of -300 s. Certain

storable bipropellants can have actual Isp values of 310-320 s. A comparison of

propellant needed to execute the Pluto orbit insertion, as related to Isp, is shown in

Figure 3.3. Although bipropellants usually have better performance values, firing

39
them is a much more complicated process because of the need for exact mixing.

This possibility for error will result in more unused propellant than one would get

with a monopropeUant (6. Koepke, p 20).

Other desirable qualities of propellants include high specific gravity and a low

freezing point. A relatively high specific gravity will result in smaller tanks (12.

Sutton, pp 168-182). Our mission requires minimizing mass wherever possible, so

bulky tanks brought on by a low density propellant could jeopardize the success of

the mission. At the distance OPTIC will be from the sun, the temperature will be

extremely cold. Frozen propellant prior to the orbit insertion could also threaten

the mission's chances.

The final choice was the bipropellant consisting of nitrogen tetroxide and

hydrazine (N204/N2H4) in use with a LEROS I engine. The mixture of N204/N2H4

combines stability with performance when coupled with a LEROS 1 engine. The

LEROS 1 is capable of 500 N of thrust and an actual Isp of 316 s. The current LEROS

is configured to perform with MON3 as the oxidizer and either hydrazine or

monomethyl hydrazine as the fuel. It is a very small engine with a thrust chamber

length of 12.7 cm and a mass flow of only .162 kg/s (4. Gray, pp 2,15). Because of the

oxidizer adaption, an Isp decrease to 315 s will be used for the design.

Most of the criteria that is desired in a propellant are satisfied by the

N204/N2H4 combination. Hydrazine is storable for long periods and has been used

repeatedly for deep space probes. It is compatible with four types of stainless steels:

303, 304, 321 and 347. It has one of the higher densities of liquid fuels at 1008 kg/m3.

The major drawback of hydrazine is its high freezing point of 274.3 "K (12. Sutton,

pp 170-181). Nitrogen tetroxide also displays some great properties as an oxidizer. It

is compatible with all stainless steels so this eliminates the need for different tank

materials (10. Parcel, p 508). Like hydrazine, nitrogen tetroxide has a very high

density of 1447 kg/m 3 and a high freezing point of 361.5 "K. However, nitrogen

40
tetroxide makes up for its high freezing point by being able to "...be stored

indefinitely in sealed containers made of compatible materials." (12. Sutton, pp

171,178) The high freezing point can be overcome by strategic placement of heaters

(see Structures subsystem). Since hydrazine and nitrogen tetroxide need to be kept

at approximately the same temperature, regulation should not be a problem. This

combination is also hypergolic, so no igniters will be needed; just mixing (12.

Sutton, pp 170-181). This propellant combination is dependable and should perform

at Pluto when it is needed the most.

Because of its size, one would assume that the LEROS engine is a poor option.

However, since reliability has been the main factor stressed throughout the design

process, this makes the LEROS an excellent selection. In order to make the Av at

Pluto of 4927 m/s, it will require a long burn. There will also be trajectory control

maneuvers (TCMs) throughout the mission. At its present configuration, the

LEROS has a firing time of about 12000 s (4. Gray, p 7). If more than one engine is

used, this is more than enough time to complete the mission's Av requirements.

Along with its proven reliability, its performance values are virtually unmatched by

any other engines with storable propellants.

The pressurization technique that will be used will be used is basically none.

For propellant feed, a technique called liquid reorientation/settling will be

attempted. This technique depends on the bond numbers (Bo) of the propellants.

As long as Bo > 1, settling of the liquid will occur and the engines will be providing

the pressure. This has been used successfully in both the Saturn V and the Centaur

upper stage (2. Cramer, pp 1-5). As shown in Appendix 3.1, the minimum thrust

needed for settling is 1.45 N. A diaphragm that only lets fluid out will be at the

bottom of both tanks. There 'should always be enough propellant in the lines to

provide this small thrust. However, in a case where there is not enough fluid in the

lines, there will be two 15 N thrusters that are pressurized with small tanks that are

41
to be used for attitude control (see that section). These thrusters will be pointed in

the same direction as the engines so settling should still be able to be achieved.

Now that propellants, engines and tank materials have been chosen, they

must be sized according to the Av at Pluto and any TCMs expected throughout the

mission. Every subsystem was told to minimize mass wherever possible. The

results were excellent and are in Table 3.2. Because of my present level of education,

all Av's will be calculated using the impulsive burn approximation (see Appendix

3.1). It was assumed that 20 out of the 40 kg of hydrazine alloted to the attitude

control system would be used prior to the burn at Pluto. Another assumption was a

structural efficiency (_) of .08 (Buckmaster, 1989). This is a very conservative

estimate because there will surely be advances in materials within the next decade.

However, it is probably accurate for this problem because of the need for extra

insulation for storage of the propellants.

The next constant that needed to be set was the percentage of extra fuel

T

needed for last minute corrections and maneuvers once in orbit about Pluto. Figure

3.4 shows the amount of propellant needed against the percentage of extra fuel

included. Figure 3.4 also shows the approximate Av capability once orbit about Pluto

is achieved. This was the deciding factor. The percentage was set at 4 for the initial

calculation and the amount of propeUant added to make the orbit insertion was not

too burdensome. The fact that Charon, Pluto's moon, is relatively close, there will

probably be a need to execute some orbit corrections during the mapping. Shortly

after orbit insertion, an orbit correction that requires a Av of 47 m/s will be executed.

This leaves OPTIC with about 400 m/s of additional Av's.

42
FIGURE 3.4

Mass of Propellant needed for Plutonian Orbit

Insertionand Av Capability After Insertionas Related to % Extra Fuel

500O

4000

 000

design

1000 ' ' 0

0 2 4 6 8 10 12

% extra fuel

2OOO

_- Mp (kg)

- A v (m/s)

•1000

A problem encountered with these iterationsistheircircularity.In order to

overcome this,afteriterationswere made, certainvalues had to be fixed. Please

referto Appendix 3.1 for detailson these calculations.After the firstpropellant

mass iterations,the mass of the extra fuelwas fixed at 95 kg. Then, more propellant

needed to be added forTCMs. A TCM capacityof around 400 m/s was incorporated

into OPTIC. Because of the duration of the mission and the need for a precise

J

approach to Pluto,thishigh TCM abilityisworth the added mass.

The small mass flow of the LEROS engine makes itnecessary to have more

than one engine to perform the Av at Pluto. Multiple engines are also needed to

keep the burn time under LEROS' 13000 s rating. The shortening of time reduces

the distance over which the burn will be made and, in turn, make the impulsive

burn approximation more accurate. Figure 3.5 illustrates

43
vehide

path

Pluto

TABLE 3.3 Engine Configurations

r •

5

4

3

2

1

Burn

3168.57

3960.71

5280.95

7921.43

15#_82.85

Load

.16

.13

.10

.06

.03

FIGURE 3.5 Physical Length of Burn Time

E

_deg)

5.28

6.56

8.73

13.02

25.36

Total angle

_deg)

15.84

19.68

26.19

39.06

76.08

FIGURE 3.6 Tank Sizing

Pumps, heaters and pipes

N2H 4

--@-

N204

I--

1.76m

.bm

.15 m

T

.67 m

.1 m

44
the proposed path and angle rE) that OPTIC will travel past periapse of its orbit about

Pluto. Table 3.3 displays results for different engine configurations. Using the

impulsive burn approximation it is assumed that the burn will take place at

periapse. In this case, the burn will be spread to both sides of the periapse point with

a total angle roughly 3 times E. Five engines will be used and will impart a load of

only .16 g's on OPTIC. The 5 engine configuration leaves room for failure. Three to

four days prior to the orbit insertion, the engines will be test-fired and if one or

more fails, the firing system will be reprogrammed for the number of engines that

work (see Command and Control sec.).

The sizing of the tanks had a limit of a 1 m diameter cylinder. The tanks are

half-spheres connected by cylinders that are part of the outer cylinder (see figure 3.6).

The half-spheres, along with a .15 m spacing between the two tanks, leave plenty of

room for wiring, heaters and piping. A thickness of 1 cm was assumed and this

thickness will guard against any micro-meteorites (10. Parcel, p 504). A redundant

piping system will be used in case of any failures. The hydrazine tank will supply

the attitude control thrusters with propellant. Figure 3.7 illustrates the redundancy

and the different supply routes. Both the attitude control thrusters and the main

engines will have 2 feeds. The final dimensions of the tank, excluding engines, is a

cylinder 1 m in diameter and 4.68 m in length.

OFTIC's final configuration is set. OPTICs propulsion system was only

designed to handle the Av at Pluto, TCMs throughout the mission and supply the

attitude control thrusters with propellant. An expendable stage is needed to supply

the Av of 7871 m/s out of LEO in order to send OPTIC on the correct

45
HGURE 3.7 Propellant Piping

System

lJ

Two 15 Eight .5

N N

Hydrazine

i

Five 500 N thrusters

trajectory to Pluto. The stage that can provide this level of performance is the initial

Space Transfer Vehicle (ISTV). The ISTV is a liquid oxygen/liquid hydrogen system

that will be available by the late 1990's. It will carry a propellant load of 28182 kg and

provide an Isp of 482 s (2. Cramer, pp 1-3). The ISTV is 11.81 m long and has a

diameter of 4.27 m. It will be designed for compatibility with the current Space

Shuttle or the proposed Shuttle-C and will give OPTIC a Av of up to 7993.35 m/s out

of LEO (see Appendix 3.1) (2. Cramer, pp 1-3).

The total mass of the ISTV and OPTIC is 34544.937 kg with a total length of

19.39 m. These values are well within both the Shuttle-C's lift capability of 45359 kg

and length of 24.7 m. Shuttle-C will be ready for launch capability in the mid-1990's

and it will be the launch vehicle that will deliver OPTIC and the ISTV to LEO (7.

Kolcum, pp 123-125).

3.3 CONCLUSION

46
OPTICs mission criteriapresented a number of problems to be addressed.

Reliability was the main feature in selecting the various components from the RTGs

to the LEROS engine. Even with the amount that reliability was stressed, there will

stiU be some potential problem areas. One of these areas is the potential for the

failure of components in the power subsystem. Twenty years is a very long time to

not have an electrical component failure. The redundant, parallel circuit will

hopefully prevent a small failure from becoming catastrophic. Another problem

that could be encountered ispropellant feed once OPTIC isorbitingPluto. There

will only be a small amount of propelIant floating in the tanks and settling that

small amount to feed the attitude control thrusters will be difficult. This design

aspect should be analyzed more if the design advances beyond the preliminary stage.

The final mass budget for the power and propulsion subsystem is in Table 3.4. Refer

to Table 3.2 for the various subsystems' mass budget. Overall, the power and

propulsion subsystem should provide a reliable support for OPTIC's other systems

and successfully execute an orbit insertion into Plutonian space.

TABLE 3.4

Final Mass Budget for

Power and Propulsion Subsystem

Cnrnr_on@n_

Power regulator

Power control un/t

Power distribution

RTGs

Tanks and pumps

Hydrazine

Nitrogen Tetroxide

ISTV stage

4.213

6.24

4.86

25.582

271.361

1751.473

1369.178

30870.0

47
APPENDIX 3.1

EQUATIONS AND CALCULATIONS

Calculation of bond numbers:

Bo _ar_tP

for N2H4:

B o - 2538.18a

r t = .48 m

p --I008-_-, o = .0915-k--E

m3 s2

for Bo > 1 --_ a > 3.94x10 -4

To calculate the minimum T required: T > aMsc = (3.94x10-4)(3684.937) = 1.45 N

for N204: p---1447kg'°=m3 .0275_

Bo ffi 12123.23a

Following same procedure, T > .30 N.

for Bo > 1 _ a > 8.25x10 -5

.'. rain. T required for settling = 1.45 N

Initial propellant ma_$ iterations:

impulsive burn approx, x Av = goISpxln(MSI+MCC+MAC+MOS+Mps+MTp+MEF+Mp)

MsI+MCC+MAC +MOS +MPS +MTP +MEF

where: go -- 9.81 _, Isp - 315 s, MSI = 56.93 kg, MCC = 80.0 kg, MOS = 67.0 kg, Av = 4927 ms

MAC -- 48.1 kg (20 kg of N2H 4 used before Pluto burn), Mps = 40.895 kg

assume structural efficiency =-e = .08- MTp

MTp+Mp

MTp= MP

11.5

set MEF = .04Mp

solving for Mp = 2292.244 kg --_ MEF = 91.69 kg, set MEF = 95.0 kg

48
APPENDIX 3.1(cont)

TCM determination:

want Av = I00 s_-

use impulsive: Av = goISpxln(MPt°t+MSI+MCC+MAC+MOS+MPS+MTP+MPTCI_)

MPtot+MsI+MCC+MAC+MOS+MPS+MTP

where MPtot = Mp(from above) + 95 + 20(attitudecontrol)kg

MPtot+MPTcM

MTp = 11.5

solve for MPTCM = 95.968 kg. add thisto find new Mp at Pluto

new Mp@Plut o = 2377.251 kg

repeat above steps up to Av = 400 _ using previous Mp's

Final values: MPTCM - 439.109 kg, MTp = 271.361 kg

Mp@Plut o = 2566.542 kg, MEF ---95.0 kg

MSC = MsI+MCC+MAC+MOS+Mps+MTp÷MPTcM+Mp@Pluto+MEF = 3684.937 kg

Because of the circularity of the problem, the Av was only a figure to base calculations

on. Now that masses are final, calculate final TCM ability.

M$C ) = 392.09 rn

impulsive approx. AVTCM = g°Ispxln(MS C . MPTC M s

Av capability once around Pluto:

MsC@Pluto - MSC - MPTCM - Mp@Plut o - 20 kg N2H 4 (used for attitude control)

= 659.286 kg (includes another 20 kg N2H 4 for attitude control)

Mp after burn = MEF --_ Av = _,Ipxln( MsC@Plut° ) = 480.82 m--

v,, MsC@Pluto. ME F s

AVreq. = 47 s_ .'. we have an extra Av capability of _ 433 m

49
APPENIX 3.1 (cont)

Burn time at Pluto:

for 5 engines:

!

acceleration(g's):

tb -- MI_@Plut° -- _ = 3168.57 s

#eng.(fit) 5(.162)

a _._.Ay_=.__49_2Z_= 1.55 m ..# g's =__a_= .16

tb 3168.57 s2 9.81

[

angle traveledpast periapse: tb - V (_-)

x(E - esin E)

where a = 9750 km, g = 663.5622.kin3

s2

solve for E using above values: E ffi 5.28 °

total angle traveled during burn - 3E - 15.84 °

repeat down to I engine, results in Table 3.4

Tank sizing:

v

Assumptions: cylindrical shell with diam. = 1 m, 1 cm thick, elongated

f,

spheres for tanks.

= 4rr_3 = .463 m3

r t=.48m Vsphere 3 t

vol. N20 4 = MN204 = 1369.178 = .946 m3 --_ vol. N204 - Vsphere = Vcyl. = .483 = 7rr2h

PN204 1447

solve for h _ hN204 = .67 m

vol. N2H 4 - MN2H4 _ 1.738 m3, same calc. as above: hN2H4 = 1.76 m

PN2H4

refer to Figure 3.6

Av out of LEO:

impulsive approx., Av = goISpxln(MISTV+Mp+Msc)

MISTV+MSC

where: MISTV = 2688 kg, Mp = 28182 kg, MSC = 3684.937 kg,

Av = 7993.35 m

S

Avpossible > Avneeded by = 122 ms

Isp = 482 s

50
3.4 BIBLIOGRAPHY

lo

o

o

Q

Bennett, Gary L. et al. "Development and Use of Nuclear Power Sources for

Space Applications," The Tournal of the Astronautical Sciences. Vol XXIX,

No. 4, Oct-Dec, 1981.

Cramer, J. and N. Brown. "Cryogenic Fluid Management Technology

Requirements for the Space Transfer Vehicle," AIAA. July, 1989, p. 9.

Giampeoli, R. "Galileo Spacecraft Mass Report and Equipment List," Issue #36.

May, 1988.

Gray, C. and R.S. Wood. "Development of a 500N Bipropellant Apogee Engine

Capable of using Hydrazine and Monomethyl Hydrazine," AIAA. July, 1989,

p. 17.

5. Koepke, Andrew et al. "Electrical Power Subsystem," AAE 241 class notes.

Spring 1990.

6. Koepke, Andrew et al. "Propulsion Subsystem," AAE 241 class notes. Spring

1990.

Sutton, George P.

Inc., 1986, p 361.

7. Kolcum, Edward H. "Decision on Shuttle-C Expected After Internal NASA

Review," Aviation Week and Space Technology. June 19, 1989, pp 123-125.

8. Mondt, J.F. and B.J. Rock. "rhermal-To-Electric Conversion."

9. Nichols, Robert G. "Showdown at Pad 39-B," AcL Astra. Nov 1989, pp 8-15.

10. Parcel, R.W. and J.B. Neiger. "Metallic and Ceramic Materials For Structural

Applications," _pace Materials Handbook. 2 ed.

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

12. Rocket Propulsion Elements, 5th ed., John Wiley & Sons,

(Buckmaster, 1989) e was taken from AAE 334 class notes, Fall 1989.

51
4.0 ATTITUDE AND ARTICULATION CONTROL

INTRODUCTION

The primary objectivesof the AACS are discussed below. They include the

stabilizationof flightduring allphases of the mission, orientationof the craftfor

communications, trajectorycontrol maneuvers (TCMs), data collection,and

determination of the relativeposition of the spacecraft.

This sectionwas researched and prepared by David Mark Robinson.

4.1 SYSTEM OBJECTIVES

As stated in the request for proposal, the objectiveof thisprojectis to develop

a conceptual design for a spacecraftto study Plutonian space. The optimization of

performance, weight and cost are very important. Our choice of an orbiting

spacecrafthas made the optimization of weight substantiallymore demanding with

every kilogram of additionaldry mass adding 8.4 kg of fuel. The attitudeand

articulationcontrol system (AACS) designed for OPTIC isbelieved to provide

versatilityand reliabilityat a relativelylow weight.

4.2 A YrtTU'DE DETERMINATION

The attitude of this spacecraft will be determined with the use of a fixed head

star tracker, a two-axis sun sensor/horizon sensor, two secondary sun sensors, and a

state of the art fiber optic rotation sensor. This combination provides reliability,

simplicity, and again relatively low weight.

52
The primary sun sensor incorporated for use of this craft is a two axis mask

sun detector providing attitude determination about two axis and horizon sensing

capability for the Plutonian orbit. A sensor of this type has low power requirements

and its light weight makes it an excellent choice for this mission. This device will

operate as a sun sensor in the cruise mode and when near enough to Pluto it will

switch to Plutonian acquired cruise. When OPTIC begins to orbit Pluto this device

will operate as a horizon sensor. The primary sun sensor and one of the secondary

sensors will be located on the scan platform with the primary sensor pointing the

same direction as the cameras and the secondary on an adjacent side (figure 4.1).

Figure 4.1"Sun sensor location on scan platform

___1 J

_ +y-axis

y

+x-axis

,Scientific instrumentation

Primary sun sensor

Secondary sun sensor

View of science platform from x-axis

53
The additional sun sensor will be located on the electronics bus opposite the

boom of the scan platform (figure 4.2). The secondary sun sensors will primarily be

used during sun acquisition and as a backup for the primary sun sensor. Mass

estimates for these sun sensors are based on sensor found in NASA documentation

(4. Giampeoli).

A fiber optic rotation sensing (FOILS) system has been chosen for inertial

reference (3. Draper, p.14). This technology is not yet available but is expected to be

by the mission dead line of 1999. The characteristics of the FORS that make it

advantageous are many. First of all the system provides three axis rotation sensing

with full redundancy which is very important in an attitude determination system.

Its residual drift rate, power requirement, and weight are low relative to present

inertial reference hardware. This system also has an extended lifetime expectancy

due to its lack of moving parts and the fact that it contains no short term wear-our

54
FIGURE 4.2

Location of AACS equipment on electronics bus.

I I

m,

i

i-x direction

[]

m

[]

Scale

.1meter

StarTracker

Sun sensor

FOPS

In order to conserve energy during the long cruise to Pluto the FOPS willbe

turned off when the craftisin cruise mode. Itwill be reactivated,however, before

any type of maneuver or ifsun acquisitionislostfor any reason. During data

collectionand transmission at Pluto the FOPS will be activated to ensure pointing

accuracy. With the FORS, pointing accuraciesas low as 0.1 degrees are realistic(5.

Hansen, p.111). This satisfiesthe requirements for antenna pointing during data

transmission to Earth. The FOPS will be located on the electronicsbus as near to the

rollaxis as possible (figure4.2).Table 4.1 listsome of the specificationsof the FOPS

system.

55
Fiber Optic Potation Sensor (FORS)

Residual Drift Rate Noise Angular resolution Availability Powe" Mass Volume

Rate deg/hr des/sec arc-sec yrs watts kg in3

2E-4 IE_ 0.00S 10 <10 10 1000

TABLE 4.1

(Draper, p.14)

The star trackerchosen for our spacecraft,the Canopus/Tracker, was chosen

primarily because of itsproven performance, lightweight (4.3kg),and low power

requirements (4.5w). Itwill be located on the electronicsbus where itwill have a

relativelyunobstructed view of space (figure4.2).This trackerissignificantlylighter

than the newer advanced star trackers (ASTROS), 23.7 kg lighter, and with a fuel

addition of approximately 8.4 kg for every I kg of dry mass this is a savings of about

200 kg of fuel. The CBS tube used by this tracker is no longer made so slight design

changes may be necessary if one can not be located. It is felt that the favorable

attributes of this particular device warrant the modifications.

As further backup for the attitude control system the camera chosen for this

mission by the science officer can image star fields for navigation.

MODES OF OPERATION

All data gathered by the attitude determination hardware will be channelled

to the spacecrafts command computer where it will be processed and the appropriate

steps will be executed depending on the mode of operation currently in effect. There

will be several modes of operation which are discussed below.

After OPTIC has been released from the bay of Shuttle C an autonomous

control system will stabilize and orient the entire system for the initial delta-V. The

56
booms and primary antenna will not be deployed until after this maneuver

considering the excessive g-loading involved. After separation from the launch

vehicle, sun acquisition will follow utilizing the three sun sensors. The FOILS will

be enabled to ensure that rotation rates are not excessive for boom deployment.

Once OPTIC has been stabilized, equipment will be deployed and sun-acquired cruise

will follow.

Since the secondary antenna is powerful enough for communication with

Earth until OPTIC reaches Jupiter the spacecraft will travel in sun acquired cruise

with the main dish poin_-_g in the direction of flight until after the necessary TCMs

are made on the approach to Jupiter. After the Jupiter flyby it will be necessary to

turn the spacecraft 180 degrees and point the larger antenna back to Earth for

communication purposes. For any TCMs thereafter the craft will have to be rotated

back and forth to insure communication with Earth for the majority of the flight. In

the sun-acquired cruise mode the FORS will be turned off to conserve power

leaving the responsibility of attitude control to the sun sensors and the star tracker.

FORS will only be used in the cruise mode when TCMs and for a short period every

24 hours to monitor the roll rate. No rotation about the pitch or yaw axes will be

allowed but the craft will be allowed the roll within deadband constraints. The

deadband roll rate will be 5 rev/min or 0.5236 rad/sec. This is to assure that the craft

does not spin out of control on its roll axis. If FOILS senses a greater rotation rate

than the deadband the roll axis thrusters will be activated to reduce the roll rate. To

reduce the role rate from 0.5236 rad/sec to a full stop when the craft is fully loaded

with fuel it will take a steady burn for 2655.2 sec of the four thrusters opposing the

rotation (see Appendix 4.1). If anything is encountered during the cruise mode that

warrants investigation the FORS will be turned on for enhanced pointing accuracy

and control. When the data has been taken and transmitted back to Earth the cruise

mode will resume (FORS off).

57
The scan mode designed for this spacecraft will be initiated on the approach to

any entity that the craft has been instructed to observe or that is determined to

worthy of investigation. The system will confirm approach parameters and make

necessary corrections in the trajectory. When within scanning distance the system

will respond to preprogrammed commands and/or to stimuli imposed by

unforeseen phenomena. Upon approach to Pluto the scan mode will measure

approach parameters and any necessary TCMs will be executed. Once of the correct

trajectory the spacecraft will be tuned 180 degrees for its deceleration burn (FORS

on). At Pluto the scan mode will include orbit station keeping, stability, and

orientation requirements imposed by scientific instrumentation and

communications. Specifically, this mode will include the mapping of the planet

which includes the collection of the data during closest approach to Pluto and data

transmission through the apogee of the orbit. Data transmission imposes the most

demanding pointing requirements and this is where the FORS will be very effective.

The magnetometer experiment also imposes requirements on the attitude

control system. A rotation rate of 3.15 rev/min or 0.32987 rad/sec about the roll axis

is desirable for the most accurate data collection (see Appendix 4.1). To avoid

possible problems of data transmission to Earth while the craft is spinning, all

magnetometer data will be collected and then OPTIC will be despun before data

transmission commences. The magnetometer experiment will be executed during

the last 3x10 orbit and the first 3x5 orbit. Despinning the craft assures pointing

accuracy.

THREE AXIS HYDRAZINE lET CONTROL SYSTEM

The final decision to use a gas jet system resulted primarily from the choice of

an orbiter mission since this significantly increased the size and weight of the

58
spacecraft. The amount of fuel required for this mission has made it necessary to

virtually send up a drum of fuel with the science instrumentation, power supply

and additional equipment on booms with the exception of the electronics bus.

Much consideration was involved in the final choice of a hydrazine jet system.

Also, the longevity of this mission puts the restraint of bearing lifetime on some

hardware. Hydrazine jets provide the torques necessary to maneuver a craft the size

of OPTIC with more agility than momentum wheels or control moment gyros and

utilization of an already present fuel source helped to reduce the weight of the

system significantly. There has been concern in the past over the affects of the

exhaust produced by hydrazine thrusters on scientific instrumentation but it was

not found to present a problem (7. Wertz, p.208).

The configuration of three-axis hydrazine jet attitude control system for the

proposed spacecraft is believed to provide ample maneuverability and stability with

sufficient redundancy. The hydrazine jet control system consists of eight 0.5N jets

for roll axis maneuvers and two 15N jets for pitch and yaw maneuvers. Thruster

locations are shown in figure 4.3. The four roll thrusters not visible in the diagram

are symetrically located at the opposite end of the main body of the craft near the

electronics bus (Figure 4.2). With the fuel source of the primary propulsion system

being hydrazine and nitrogen tetroxide, a monopropellant control system using

hydrazine seemed to be a good opportunity to save weight since this eliminates the

need for additional fuel tanks if the primary hydrazine fuel source can be utilized.

This also adds a margin of safety considering that if the control system fuel budget is

exceeded there should be more than enough in the primary systems fuel budget to

compensate. The concern of fuel supply is addressed in a later section. The fuel

budget of 40kg is based on estimates using a similar system described by Wertz. (7.

Wertz, p.209)

59
As mentioned earlier, the 0.5N jets will provide roll torques to the space craft.

The torque will be produced by four gas jets at a lever arm of 0.5 m providing a 2 Nm of torque. The system designed provides complete redundancy allowing up to

four of the eight thrusters to be disabled and retaining control of the craft. The 151'4

jets are mounted at the rear of the craft, the line between them being parallel to the

y-axis. They will be used for applying pitch and yaw torques( figure 4.3). By

orienting these jets such that the line between them is parallel to the yaw axis and

firing one jet a pitch torque is produced. In this position firing the top jet will

produce a negative torque or firing the bottom jet wifl produce a positive torque.

Yaw maneuvers are accomplished using a similar procedure aligning the line

between the two jets parallel to the pitch axis (figure 4.3).

The two 15N jets wifl serve the additional role on this mission of settling the

fuels in their respective tanks before major burns of the primary propulsion system.

To assure fuel supply to the 15N thrusters a lkg capacity nitrogen pressurized

hydrazine fuel source has been provided for each jet. Once the settling burn has

been completed these jets will then resume feeding off of the primary hydrazine

fuel source.

The pulse duration of the burns will range form 0.1 seconds to up to several

minutes depending on the desired maneuver. Pulse burns a more efficient

however the longer the burn (6. Sutton). Optimal pulse durations are a trade

between fuel consumption and the require speed at which the maneuver must be

executed. This optimization is beyond the scope of this proposal. Mass estimates are

based on similar hardware described in a NASA document (4. Giampeoli)

For trajectory control during burns of the primary propulsion system a thrust

vectoring device will be utilized. This should facilitate excellent control during long

duration burns and take the burden of some trajectory control maneuvers off of the

smaller hydrazine jet system, hopefully resulting in lower fuel consumption.

60
0.SN roll thrusters

FIGURE 4.3

Location of attitude control thrusters

i x-axis (pitch axis)

15N pitch/yaw thruster

© ©

z-axis

Uaxis

© ©

y-axis

(yaw axis)

4.2 ARTICULATION CONTROL

For scan platform articulation control a micro step actuator will be used (3.

Draper, p.14). This is a new device which incorporates momentum compensation

technology and will give increased pointing accuracies for the scientific

instrumentation. It will also be used during sun acquisition to give the sun sensors

on the scan platform a greater field of view.

The secondary antenna also requires articulation control. An actuator with

three axis control is needed to give it the pointing freedom required for its optimal

61
use. The more freedom this antenna has to move the less the spacecraft is required

to move for it.

The masses of these devices were not located, but estimates were made using

NASA documentation (4. Giampeoli).

POSSIBLE PROBLEMS

Without the use of a bladder or a pressurization system for the hydrazine fuel

tank, fuel supply could pose a problem for the attitude control system. The addition

of the nitrogen pressurized fuel sources for the 15N thrusters assures them of fuel,

however, the 0.5N roll jets are not equipped with these pressurized tanks. This was

primarily to reduce weight. A remedy for this problem could be a short duration

pulse, approximately 0.1 sec, of the 15N jets supplied by the pressurized tanks to

settle the fuel in the main tank. After this pulse all thrusters would then resume

feeding off of the primary hydrazine source. Also, to remedy this problem a one

way valve has been used as the feed valve for all propulsion systems. This locks

fuel in the lines for assured short duration burns.

Another consideration that was overlooked until lat in the design was the

need for multiple sun sensors and their placement. Originally only one primary

and one secondary sensor were thought necessary. However, further research

revealed the possible necessity of more sun sensors to assure sun acquisition. The

final configuration decided upon consists of two sensors on the scan platform that

wiU be able to scan space with the sweeping motion of the platform and a third

located on the opposite side of the spacecraft in the electronics bus out of the field of

view of the others (figure 4.2). This configuration may be inconsistent with other

information in this proposal but it is the recommended configuration.

62
As can be noted in Appendix 4.2 the roll axis moment of inertia (Lzz) changes

from the smallest to the largest of the three axial inerfias sometime during the

deceleration burn at Pluto. It would be likely that sometime during the burn that

Izz will be the intermediate value rendering it the unstable axis. Thrust vectoring

and the control system should retain stability but this is still an area of concern.

4.4 AACS MASS SUMMARY

COMPONENT

i

Fuel Budget

Roll Thrusters (4 @ 0.5 k_ each)

Pitch/Yaw Thrusters (2 @ 1 kg each)

i

Pressurized Fuel Tanks (2 @ 0.1 k_ each)

Star Tracker (Canopus/Tracker)

Sun Sensors

i

FORS

Thrust Vectoring Assembly

Secondary" Antenna Actuator

Scan Platform Actuator

I

Total:

MASS (k$)

42

4

2

0.2

4.3

~ 0.2

10

3

(Possibleunderestimation)

3.1

1.5

70.3 ks

63
APPENDIX 4.1

Calculation of pulse duration for specific roll axis maneuvers.

Torque = Force x Lever arm -- Moment of inertia x Angular acceleration

'C= l:'I., = I(Z

(X= FL/I

integrate once d(x/dt = co = FLt/I + const. (const. = COo)

integrate again dco/dt = theta = FLt2/21 + COot+ Const.

Assume all constants to be zero for these calculations.

F = Sum of forces of all thrusters acting (N)

L = Lever arm (m)

I = Moment of Inertia (kg m 2)

co = angular velocity (rad/sec)

a = angular acceleration (rad/sec 2)

For worst case of deceleration from deadband (5 rev/min) to full stop with full

tanks.

t = coI/FL - (.5236 rad/sec)(5071.05 kg m2)/4(0.5N)(0.5m)

t = 2655.202 sec = 44.25 min.

For spin requirement of the magnetometer experiment of 3.15 rev/min

t - (0.329867 rad / sec)(4378.506 kg m2) / 4(0.5N)(0.5m)

t = 1444.325 sec = 24.072 min.

Pitch and yaw burn duration for 180 degree maneuver

Before deceleration burn: t = (2_/FL)0.5 = ((270(9859.6)/(15N)(0.Sm))0.5

t = 90.88 sec = 1.515 min

After deceleration burn: t = ((2_)(3835.062)/(15N)(0.5m))0-5

t = 56.682 sec = 0.9447 rain.

64
4.5 BIBLIOGRAPHY

lo

2.

o

o

o

1

California Institute of Technology - Jet Propulsion Laboratory. Mariner - Vcnu_

1962, Washington D.C. (1965).

Chobotov,V.A. Spacecraft Attitude Dynamics and Control, Kinkos notes (Spring

1988).

Draper, tLF. '_l_e Mariner Mark II Program," Jet Propulsion Laboratory,

Pasadena, Calif.

Giampeoli, tL "Galileo Spacecraft Mass Report and Equipment List," Issue #36,

(May 1988).

Hansen,D.M. and Kliore, A.J. "Ka-Band (32 GHz) Benefits to Planned

Missions,"(Oct - Dec 1986).

Sutton, G. Rocket Propulsion Elements, 5th ed. John Wiley & Sons: New York

(1986).

7. Wertz, J. Spacecraft Attitude Determination and Control, Kluwer Academic

65
5.0 STRUCTURES

INTRODUCTION

This section is a discussion of the overall OPTIC structure, the materials used,

and the final configuration of the spacecraft.

This data presented in this subsection was written and investigated by Mark

James Endre.

5.1 STRUCTURES SUBSYSTEM OVERVIEW

As outlined in the Request for Proposal (RFP), once the OPTIC mission was

decided upon, each subsystem must meet certain requirements. Since the launch

date is between the year 2000 and 2010, the spacecraft must use materials and

techniques available prior to the year 2000. These materials and subsystem

components should be "off the shelf" whenever available, be reliable, easy to

operate, and be relatively inexpensive.

An important factor in the design of the spacecraft configuration was the fact

that there were two semi-conflicting requirements in the RFP. They were that once

the orbiter was decided upon, nothing in the design could preclude the spacecraft

from performing several different missions.

The main concern of the structure subsystem was to make sure that the craft

could survive the launch sequence, the space environment, and be able to complete

the outlined mission. Of the three requirements, the last was the most difficult.

Since the mission type was an orbiter, several limitations were placed upon the craft

itself. The most important of these was the fact that all subsystem masses had to be

66
minimized. The design also required further optimization to meet the low mass

requirements.

5.2 STRUCTURAL CONSIDERATIONS

There are many different structural considerations for any interplanetary

mission. It was decided that for any mission in Plutonian space, an orbiter, if

feasible, would be the most cost effective. However, the decision of using an orbiter

created many technical hurdles that had to be overcome to prove feasibility. The

main problem facing the structure subsystem was the reduction of the spacecraft

mass.

MASS REDUCTION

The problem of reducing the mass as much as possible was attacked in

various ways. Three prominent methods of reducing mass were considered. They

were: the use of new extremely strong materials; using necessary components in a

dual purpose capacity; and optimizing design techniques.

The advances in material sciences tends to suggest that the newly improved

materials technique would be used. More specifically the use of a "smart" material

that knows when to change properties could be very useful (4. Of Material, pg. 22).

These smart materials are materials with an electrorheological fluid embedded in

them. These fluids change viscosity almost instantly when a low-amperage, high

voltage current is passed through them. These materials could be used to increase

the strength of other materials and be used as a low weight, active dampening

system. Although smart magnesium and aluminum metals have been fabricated,

most research has been dealing with composites. Unfortunately, these materials

67
along with the other advanced materials, have not had rigorous long term testing

that would prove their space readiness. Since these new materials have not been

space proven, it is possible that they could fail over the course of the mission.

Consequently it was decided that this option conflicted with the RFP requirements

of reliability.

The choice of a structural material was based on material properties such as

strength to density ratios, operating temperatures, and properties of the material in a

space environment. Aluminum is inexpensive, does not suffer large radiation

effects, does not sublime, and would be able to be used far below temperatures at

which it deteriorates in vacuum (5. Parcel, pg. 498, 2. Ashby, pg. 14). Since

aluminum has been repeatedly proven to be spaceworthy, it was chosen as the

material to be used in the electronics bus, the science platform, and in the

micrometeorite shield.

STRUCTURAL LAYOUT

The fact that the mission type was an orbiter created a necessity for a large

amount of fuel, and subsequently a large fuel tank. To reduce the overall spacecraft

mass the fuel tank was used as the main structural component. This option not

only reduces mass, but makes the spacecraft simpler and structurally more reliable.

The four extendable booms of the spacecraft are directly attached to the top

end of the fuel tank. These booms, once deployed, will take on the configuration

illustrated in Figure 5.1.

The main high gain antenna will also connect directly to the fuel tank. The

mounting will pass directly through the electronics bus coincidentally with the axis

of revolution of the fuel tank, hereafter the Z axis, as an Figure 5.2.

68
FIGURE 5.1: Top View of the OPTIC Spacecraft

Axis Y

T

RTG

3.1 m

J Platform

• _-,.3.1 m

lm

R= 2.4 m

Scale

I I

I meter

3.1m---#_

Secondary

Antenna

X Axis

.23 m

8.6 m

Magnetometer

69
YAxis

FIGURE 5.2:

ElectronicsBus

.Im

Side View of the OPTIC Spacecraft

2.4m

_m

Z Axis

t

It

I I

Main Antenna

Assembly

1.4 m

1.1 m

I

Boom mountings/Structural supports

Scale

I1 meter I

4.68 m Attitude Thrusters

(one one each side of tank)

Fuel Tank and Engine

System

70
Also centered around the antenna mounting is the electronics bus. This

component is stationed on top of the extendable boom mountings along with other

additional structural supports, and several vibrational dampeners. These

dampeners are mounted on top of the structural supports and protect the electronics

from vibrational damage.

Since the power source for the mission emits radiation, it must be stationed

far enough away from the rest of the instruments so that it will not

interfere with their performance. Using the Galileo probe as a guide, the RTG was

mounted on a five meter extendable boom; Once extended this boom was

designated as the positive Y direction in the spacecraft internal coordinate system.

See Figure 5.1.

The magnetometer is also highly sensitive to the emissions of other various

components on the spacecraft, particularly the RTG. For this reason the

magnetometer is also mounted on an extendable boom, again mimicking the

Galileo probe. The magnetometer is mounted on a ten meter boom orientated 180

degrees away from the RTG boom.

This left the science platform and secondary antenna mounted on five meter

booms that are perpendicular to the Y axis. The direction toward the secondary

antenna was arbitrarily designated as the positive X direction. See Figure 5.1.

The science platform itself has four instruments mounted on it. The only

requirement of these instruments is that they have unimpeded fields of view. All

four components are mounted on the underside of a meter square A1 2024 plate.

The three optical scanners are mounted such that they are flush with the negative X

edge of the plate. In this configuration they point away from the spacecraft when

the science platform is in the neutral position, as depicted in Figure 5.1. The last

instrument is a sun sensor. It is mounted flush with the positive Y side of the

platform, as pictured in Figure 5.3.

71
The science platform itself is mounted with actuators that allow it to rotate

independently of the spacecraft. This reduces the amount of spacecraft pointing, and

therefor reduces the amount of required fuel and total mass of the mission.

FIGURE 5.3 The Science Platform

Science Platform Viewed from Below

IILIllIIlllILIllIIIIIIIIllIIIIIIIlli

m

YAXIS

y

X

AXIS

ide Angle Camera .7m x.2m

_arrow Angle Camera .95m x.3m

_Ultra-violet Spectrometer.35m x .19m

[_Sun Sensor .lm x .lm x .05m

Science Platform Viewed from the Side

lm

.95 m

r "';

The secondary antenna is also mounted on actuators that allow independent

pointing. This function will be used in the mapping portion of the mission.

The fact that these two independently movable systems are on opposite ends

of the spacecraft allows a reduction of spacecraft maneuvering, both alone and in

conjunction. If for instance when one system is not being used, that system can

?2
move in such a way that it can negate or greatly reduce the torques applied to the

spacecraft by the system in use.

In the event of a failure of the pointing system, the pointing of the

instruments can be obtained by spacecraft orientation. If only one system fails,

spacecraft orientation will be used for pointing that system, while the operational

system will point itself. These pointing functions would be controlled by the on

board artificial intelligence.

The five main engines of the spacecraft are placed as in Figure 5.4. The

placement of the attitude control thrusters are also depicted in Figures 5.2 and 5.4.

Attitude

Thrusters

FIGURE 5.4: Engine

Configuration

O O

Main

Engines

MICROMETEOR1TE PROTECTION

An important factor of the space environment is the presence of

micrometeorites. Micrometeorite damage over the course of the mission could

conceivably cause enough damage to a spacecraft to stop its operation. To prevent

this damage to the sensitive electronics controlling the spacecraft and

communications, a three stage defense was designed.

73
The first stage defense is the electronics bus itself. The electronics of the

spacecraft are mounted inside the electronics bus, and the bus protects the

equipment from the space environment. From Figure 5.5 an electronics bus with a

2.2 m 2 area on a 25 year mission has approximately a 95% chance of not berg

punctured by a meteoroid if it is 2 mm thick (5. Parcel, pg. 503).

FIGURE 5.5

10!I

I" P 1°1=0"9999 -'_ : ' ' I IJ

I" i J/J

"E l- PIo)=0.9900.-_\ _ I / I ./1

=_ L,,o,:o, ooj

,_-2J , I,_if lJtlI IIa I I" J ,zll, _.I i ll_l ,o' _' 1o' 1o' _ lo'° _" ,o'2

AT (QZ.**¢)

FIGURE 5.5: _neet Thickness of AI as s _netton of

the Suz_ace _ea-LLfeltme Product Requtz_l for Ya_ous Probabilities of No

• Meteoroid Puncture (lq_. 5, p. 503).

The second stage is a buffer system, or more commonly a micrometeorite

shield. This shield is 1 mm thick and for maximum effectiveness is mounted 1.2

mm outside of the electronics bus. This shield covers the side and top of the

electronics bus (5. Parcel, pg. 505).

The last stage of defense is a thermal, electrostatic and micrometeorite

protection blanket similar to the one used in the Galileo probe (3. J'PL). This blanket

is inside the electronics bus and covers the electronic components.

74
THERMAL CONTROL

There were two available options for the thermal control of the spacecraft.

The first option, passive control, requires a complete knowledge of the

environment. Although passive control is generally a lower mass solution, the

knowledge of the environment was not complete enough for this option to be used.

For this reason active control is used for thermal control of the spacecraft (6. Vajta,

p 99). A full thermal analysis however is beyond the scope of this course. Therefore

thermal control is regulated with six single watt heaters in the electronics bus and

four single watt heaters in the fuel tank (7. Yeates, pg. 112).

ELECTRONICS PLACEMENT

The placement of the electronics components in the electronics bus is based

on the size of the component, heater placement, and an attempt to balance the mass

of the science platform. See Figure 5.7 for an illustration of the electronics bus

layout.

RADIATION EFFECTS

The planned trajectory at first raised a possibility of radiation contamination.

However, since the gravity assist maneuver is a great distance from Jupiter,

radiation exposure and contamination of the spacecraft is not a problem. Therefore

no additional radiation protection is necessary (1. Andrew, pg. 52).

75
FIGURE 5.7 The Electronics Bus and Electronics Layout

i

I 1

3

11

Scale

.1 meter

11 11

14

11

i

Isl

11

2

Unit #

1

2

3

4

5

6

7

10

11

12

14

15

16

17

TABLE 5.3: Electronics Compononent Masses

Component Name

Computer

Computer

Reciever System

Down Converter

Convulution Coder

Main Telemetry Modulation Unit

MTMU Backup

Colnm_I_d Detector Unit

(_DU Backup

TWTA Subsystem Assembly

Heaters

Star Tracker

Inertial Reference System

Radar Imae'iO_ Electronics

Pgwer Rcgula_gr

Power Control

Power Distribution Unit

Mass (Kg)

6

4

20

2.4

.6

ill

2.4

2.4

2.5

2.4

5.9

6@1

4.3

10

4.21

6.24

4.86

76
MASS AND INERTIA CONFIGURATION

Using masses, physical locations, generalizations, idealizations, and

dimensions of the other subsystem components, an approximate center of mass and

approximate inertia matrix were calculated for the spacecraft. The idealizations and

generalizations used were that components had uniform densities and conformed

to simple symmetric shapes. For this preliminary design these methods should

have resulted in values reasonably close to the actual values. Structure subsystem

components and masses are listed in Table 5.1, while all subsystem and component

masses are in Table 5.2. The spacecraft inertia matrix at the start of the mission and

upon arrival at Pluto are listed in Figure 5.6. These are based on a coordinate system

at the center of mass with the same orientation of the coordinate system in Figure

5.1 and 5.2.

TABLE 5.1: Structure Subsystem Masses

Component Name

Magnetometer Boom

Science Platform Boom

Secondary Antenna Boom

RTG Boom

Mass (kS)

8.81

4.4

4.4

2.8

Electronics Bus 12.17

Micrometeorite Shield 3.87

Protection Blanket 1.55

Science Platform 10

Main Antenna Mounting

Heaters

Vibration Dampeners

StructuralSupports

TotalMass

5

10

2

2

67

77
FIGURE 5.6: Inertias

Inertia Matrix With Full Fuel Load

9859.6 41.35 607.82

41.35 10268.34 -122.72

60,7.82 -122.72 5071.06

Inertia Matrix at Pluto Arrival

]

3835.06 21.90 397.61

21.90 4146.59 -82.22

397.61 -82.22 4378.50 ]

78
APPENDIX 5.1: Equations

Skirt Thickness Equation:

(Ref. 5, pp. 501-502)

t_

t = vehicle skin thickness (cm)

A ---exposed surface area (m 2)

= exposure time (sec)

o¢= 3.3 x 10"15

_ = 1.34

p(o) = probability of no

punctures

1

Shield Spacing Equation:

s = 2 tt tb Pb

t t = thickness of main plate (bus) (cm)

t b = thickness of shield (cm)

Pl_ = density of shield (gm/cm 3 )

79
5.3 BIBLIOGRAPHY

.

e

e

4.

1

Andrew, Alan et. al.. "Penetrating Radiation," Space Materials Handbook,

Lockheed Aircraft Corporation: Sur_y'vale California. (1965).

Ashby, Michael F., Jones, David R.H.. Engineering Materials 1 An Introduction to

their Properties and Applications. Pergamon Press: New York. (1988).

Jet Propulsion Laboratory, Interoffice Memo, Div 35 Mass Division. Sept. 1,1978.

"Of Material Interest," Metal Progress, Advanced Materials & Processes Inc.

(9/88).

Parcel R.W., Nelger J.B.. "Metallic and Ceramic Material for Structural

Applications," Space Materials Handbook, Lockheed Aircraft Corporation:

Sunnyvale California. (1965).

6. Vajta T.F.. "Thermal Control Materials," Space Materials Handbook, Lockheed

Aircraft Corporation: Sunnyvale California. (1965).

!

7. Yeates, C.M., et.al.. Galileo: Exploration of ]'upiter's System, NASA Scientific

Branch and Technical Information Branch: Washington D.C.. (1985).

80
6.0 COMMAND, CONTROL AND

COMMUNICATION

I1WTRODUCTION

This section describes and explains the command, control,and

communication of OPTIC. This subsystem involves data management,

transmission, and onboard computing.

This section was compiled and written by Eric W. Summers.

6.1 SYSTEM OBJECTIVES

Command, control, and communication is responsible for the well being and

good management of the spacecraft. It is also responsible for the proper

management of all data generated by the spacecraft. This data can be science data,

engineering data, or imaging data.

Science data comes from measurements made by the science

instrumentation. Engineering data is data about the spacecraft's health. Imaging

data is pictures or radar images of the planetary surface. All of these types of data are

called telemetry. This is the most important type of spacecraft communication.

There are two other types of communication: command and tracking. Command is

communication from Earth that tells the spacecraft what to do. It is generally a

rather high quality signal so that the spacecraft doesn't get an incorrect instruction.

There are numerous safeguards, such as parity checks and checkwords, to insure

that this doesn't happen. Tracking yields information about spacecraft velocity,

position, and the interstellar medium. If the signal happens to pass through an

atmosphere then atmospheric composition can also be determined.

The command, control, and communication subsystem must aid and assist

the mission at hand, which is to reach Pluto and to learn the most that it can;

81
however, there is also a set of guidelines which must be followed and, whenever

possible, adhered to.

6.2 RFP REQUIREMENTS

The RFP lists several general requirements that are applicable to the

command, control, and communication (C3) subsystem. These requirements are as

follows. There should be nothing that will limit the spacecraft to only a Pluto

mission. The spacecraft must last, at least, to complete the mission. All materials

and techniques must be available or be projected to be available by 1999. The

spacecraft must be as simple, reliable, practical, original, feasible, and as low cost as

possible. Off the shelf hardware should be utilized when possible. Components

should be optimized for weight and cost. Significant intermediate events, potential

problem areas, and concerns should be indentified. The only requirement that is

specifically directed at C 3 is that the latest advances in AI (Artificial Intelligence)

should be used to decrease spacecraft cost and increase spacecraft autonomy. This

i

requirement is somewhat at odds with the simplicity, low cost, and off the shelf

hardware requirements. A compromise will have to be reached.

6.3 COMMUNICATIONS

The communication system is broken down into four basic parts: antennas,

frequencies, telemetry, and receiving stations. All of these parts are interdependent

and require optimization.

There will be two antennas onboard the spacecraft: a main high gain antenna

(HGA) and a secondary low gain antenna (LGA). Both antennas are of the new

collapsible mesh design. This type of antenna will be used on both the Galileo and

82
Cassini missions; therefore,the design should be adequately fieldtested. The HGA

willbe the one used on the Galileo mission. This was chosen because itisthe

largestof the new mesh antennas to be designed, and itisoff the shelfhardware.

This antenna is 4.8m in diameter. The main difference is that Ka band will used as

opposed to X or S band. This is a relatively small adjustment. The feed cone will

need to be replaced to make this antenna Ka capable. The antenna surface is

optimized for X band; however, there is only a small decrease in efficiency when Ka

band is used.

The LGA will be used for spacecraft to Earth communication en route to

Jupiter. This is done to facilitate communications without having to rotate the

spacecraft 180 °. This antenna will use Ka band for communicating with Earth;

however, once Pluto orbit is achieved this antenna will be used to do radar mapping

of Pluto's surface. Radar mapping will require X and S bands. This means that the

LGA will have to be X, S, and Ka band capable. Cassini uses all three bands so this

shouldn't present any problems. The LGA will be 1.0m in diameter and three axis

gimballed so that it can be pointed in any direction in its field of view.

Ka band is the frequency of choice for a Pluto mission. The higher the

frequency the lower the transmitted power required for communication with Earth.

Ka is a higher frequency than X; therefore, less power is needed. The mass of a Ka

system is comparable, or will be by 1999, to that of an X system. Ka is currently

unproven, since Cassini is going to use Ka for communications, it will be proven in

the near future.

Ka is the newest of the frequencies to be allotted for deep space

communications, which is why it hasn't been proven yet. Ka band is 32 GHz for

spacecraft to Earth (downlink) and 34.5 GHz for Earth to spacecraft (uplink). X band

is 7.161 GHz for uplink and 8.414 GHz for downlink. Ka was chosen because it was

the last good window in the microwave spectrum. There is some concern about the

83
effects of atmospheric attenuation of a Ka signal due to water in the atmosphere.

However, there is only an 8% loss due to the atmosphere in the worst case average.

This is not a large loss. The advantage of Ka over X band far outweighs a small

atmospheric loss.

The last frequency to examine is an optical one. Optics promises a great deal:

higher bit rate, lower power, and smaller size. Currently, optics is not feasible

within the 1999 development deadline. Optical communications requires pointing

1000 times more accurately than the best X band pointing accuracies and the

development of an entirely new receiving network (9. Metscher, p.110). Such a

network is not realistic before 2010 and may have to be in orbit as opposed to being

ground based (12. Smith p.98). Considering the demands and uncertainty of an

optical communications system (8. Layland, p.123), optics is currently an

unacceptable option for the Pluto mission.

Telemetry is the most important part of the communication system. If no

data is returned then the mission is a failure. There are many things to consider

when working telemetry in smoothly with the other systems: bit rate, redundancy,

coding, data compression, when to send, how much to send, and in what order

should the data be sent.

The bit rate is based on the amount of data there is, transmission time and

computer memory available, background noise temperature, bandwidth, signal to

noise ratio (SNR), and the redundancy. The science subsystem sends data to the

computer for storage, coding, and then transmission. If the amount of data received

is greater than the amount transmitted then memory space will start to fill up. This

is not a problem until the memory reaches capacity and then data will be lost. The

amount of power available, antenna size, and the background noise is fixed.

Therefore the bandwidth and SNR must be manipulated to provide an adequate bit

rate. The amount of data that is sent can be compressed. This means that a smaller

84
bit rate can be used. The amount of time that communication with Earth is possible

is dictated by science needs and the orbit. The HGA must be pointed towards the

Earth; however, when science is taking measurements the spacecraft must be

pointed to where science wants to look. Transmission is only possible when science

does not require any particular orientation, and the spacecraft is in the right place in

its orbit. The redundancy also requires memory space. If transmitted data is stored

until confirmation from Earth is received, then a large part of the memory could be

taken up by this data. The best way around this is to have a large SN-R, so that most

of the data arrives at Earth, and a bit rate larger than that of science input. The order

that the data is sent will conform to current NASA standards.

It is evident that all of the factors are dependant on each other. This means

that optimization in an iterative process. Fortunately, most of the factors are

bounded or in some way fixed. The SNR must be greater than 10 for good

communication, transmission power has an upper bound, the maximum Ka

bandwidth is .5 GHz, and the DSN receiver is 70m. This means that the bit rate is

within a narrow range, so that optimization is easier.

The receiver system will be the Deep Space Network (DSN). There are three

major DSN sights around the globe: Goldstone California, Madrid Spain, and

Canaberra Australia. Because these sights are spread as they are there is no time that

the spacecraft cannot communicate due to the rotation of the Earth. Each of these

sights has a 70m and a 34m dish. The 70m dish is preferable for a Pluto probe

because of its larger size. DSN will be modified for Ka band by 1995 (5. Imbriale,

p.127). This modification will make the DSN more efficient, in Ka band, over a

wider range of elevation angles.

85
6.4 COMMAND AND CONTROL

Command and control iscomprised of three basic areas: command hardware

and software, the computer, and interactionswith other subsystems. The command

hardware consistsof a command detector unit and a convolution coder. Currently

these are pieces of equipment separate from the computer; however, in the future

they may become part of the software. The heart and soul of the command and

control system is the computer. A computer isbasicallya centralprocessing unit

(CPLD and memory. There will be four CPUs that will be used, 3 for processing and

1 for a backup. In case of multiple failures the computer could operate in a reduced

mode on only I CPU. The CPU that will be used is a 32 bit microprocessor being

developed by the Department of Energy (DOE) (2. GAO, p.24). The use of a

microprocessor that is currently being developed will save the program money by

utilizing off the shelf hardware. A 32 bit processor is more powerful than any

processor currently in use, or planned for use in space. This extra power will allow

the spacecraft to use higher level programming languages than are currently in use

in space. Current spacecraft are programmed in assembly language. Assembly is

difficult to program in because it is a low level language. The advantage of a higher

level language is two fold. It is cheaper and easier to program in. Low cost is an

important point in the RFP. The relative ease of programming in a higher level

language will also make the spacecraft more flexible. If the spacecraft should

experience any long delays then it could be reprogrammed for newer technology and

techniques. This would have done for Galileo had it been programmed in a higher

level language (2. GAO, p.35).

There are two choices in the selection of a higher level programming

language. The first choice is "C". C is widely used in industry and thus well known.

The other choice is ADA. ADA is a government standard language and it has been

86
proven in space by the Europeans. ADA also has a prioritizing function. This

means that when something more important comes to its attention it will stop what

it is doing and do that which is more important. These factors make ADA the

language of choice. A higher level language is required for artificial intelligence (A.I)

and expert systems. However, higher level languages are not all pros and no cons.

On the flipside, ADA will require more memory than assembly. Both ADA and AIs

will require more memory than is currently in use on spacecraft.

A computer has two different types of memory: internal and external. There

will be 8 megabytes (MB) of internal random access memory (RAM), a 660 MB hard

drive, and a 256 MB external optical drive. The 8 MB of RAM is somewhat of a

standard on commercial high end personal computers, the Macintosh TM IIcx by

Apple Inc. is a prime example. Both the 660 MB hard drive and the 256 MB optical

drive are new products of NeXT TM Inc. The selection of an optical drive over a

more conventional magnetic tape drive is easy when it is realized that an optical

system is much smaller than a magnetic system of comparable memory, and the fact

that optical disks can be written to as many times as desired without any data

dropouts or degradation. Magnetic drives cannot come close to this kind of

performance. The size of computer memory is rapidly increasing while the cost is

decreasing. This trend is what will allow the spacecraft to have an unprecedented

amount of memory. This large amount of memory will allow the spacecraft to use

the latest in AI technology, entirely backup its internal programming, and save all

transmitted data until confirmation from Earth is received. Therefore if the

internal memory is lost, the spacecraft can continue to operate from its other

memory sources. If either of the external drives should be lost the other can also act

as a backup. There should also be almost no data loss, as all data can be

retransmitted until it is correctly received. This should make the computer system

extremely reliable.

87
There will be heavy use of AI technology. The onboard AIs will monitor and

control all of the spacecraft's functions. There will be an AI to monitor the health

and status, make minor course adjustments, regulate and prioritize the onboard

systems, and control the communications system. The AI will be able to be

overridden from Earth if such an occasion should arise. This large degree of

autonomy will be necessary once Plutonian orbit insertion is achieved. There will

be a great deal to do and many new and unexpected situations will arise. Assistance

from Earth will not be quick enough as the round trip signal lag time will be about

11 hours. This is why the use of AIs will be needed. The AI technology for this is

not available to do all of these things at this time. Currently all of the above

functions ( health and status, course adjustments, etc...) can be done, but the AIs

cannot yet deal with the unknown. This may not be necessary if enough data about

the size of the Pluto/Charon system can be learned before orbital insertion. This

information could be learned by the spacecraft as it gets close to the system or

perhaps Hubble will find out the size of the system.

Command and control must also interact with the other onboard systems.

Attitude and articulation control is going to run itself off of the command

computer. This was done to cut down on mass. Science will send its data to the

command computer. Structures decides when the booms should be deployed and

the computer will do it. Should there ever be any sort of power shortage or

equipment failure then the computer will have to prior/tize and sequence

accordingly, this would be an interaction with the power and propulsion subsystem.

The command and control subsystem is responsible for the well being and good

management of the spacecraft.

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6.5 SPECIFICATIONS

The HGA will be 4.8m in diameter and transmit with 6.3 watts of power. The

LGA will be 1.0m in diameter and transmit with 2.1 watts of power. There is a DC to

RF power conversion factor of .21 (4. Hansen, p.111). This means that for the HGA

to transmit with 6.3 watts of power there must be 30 watts of power coming in.

The NeXT TM optical drive uses a 5.25" disk suspended in a polycarbonate

medium. The disk spins at 3000 to 3600 revolutions per minute (rpm). It requires

18 watts to read and 40 watts to write. There is a 92 millisecond (ms) search time

with a 18 ms search time ifthe information is within a 3 MB range. The drive can

read/write bursts at rate of 4.6 MB per second, or there isa sustained read/write of .8

MB per second. The drive must be kept between 10° and 30° Celsius.

Data willbe sent at about 388 kilobitsper second with a signal to noise ratio

(SNR) of 25. A SNR greater than 10 isrequired for good communication. A SNR of

25 was chosen so that a minimal amount of data would be loston itsway to Earth.

Data compression at 4:1 willbe used (4. Hansen, p.113). The background noise is

assumed to be 8°IC The signallosses due to the atmosphere are 8%, transmitter

pointing lossesare 11%, receiverlossesare also 11%. For calculationsand equations

see appendix 6.A.

6.6 CONCLUSIONS

There are a few significant current developments which were not included in

the design of OPTIC, but bear further investigation. The subreflector system aboard

Cassini utilizes a frequency selective surface (FSS). This would allow an antenna to

adjust itself to the desired frequency in order to optimize the transmission (5.

Imbriale, p.128). There is also the possibility that a new DSN array will be built, this

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would give a greater effectivearea received; therefore,utilizingavailablepower

better.There are also the hope thata greaterDC to RF efficiencycan be achieved.

This would also betterutilizeavailablepower.

There is a slightproblem in the communication bandwidth that must be

resolved. The current bandwidth isabout 20 kHz. Typical bandwidths are no less

than 1% of the operating frequency, or 320,000 kHz. The 20 kHz cannot be made to

be 320,000 kHz Without decreasing the SNR, an unacceptable choice,or increasing

transmitted power to 106,000 watts, also unacceptable not to mention impossible.

Since any space mission would encounter similar problems, there must be a

solution to this dilemma.

9O