SATURN ORBITER WITH DUAL PROBES MISSION: A PROGRESS REPORT" . .. ..< . .

1 C'T'5LLIUS _ _ _- - . ~ ~ Donna L. S. Pivirotto""

California Institute of Technology Jet Propulsion Laboratory Pasadena, California

Abstract

The Saturn Orbiter with Dual Probes (SO2P) mission is currently slated as the next NASA outer planets mission after Galileo. Launch is planned in the late 1980s to early 1990s. S02P will inves- tigate the atmosphere, environment, and satellites of Saturn, the Saturn ring system, and the atmas- phere and surface of Titan, Saturn's largest moon. Mission options include a variety of launch config- urations, orbiter and probe designs and configura- tions, and encaunterlsatellite tourlprobe deployment strategies. Detailed descriptions of the most attractive options are presented, includ- ing preliminary science objectives, mission requirements and constraints, mission scenarios, and definition of work required to enable the mission options. Saturn arbiter missions with Saturn and Titan probes are scientifically attrac- tive and cechnically feasible, particularly if performed synergistically with similar missions to the planets beyond Saturn.

1. Introduction

A Saturn orbiter with probes to Saturn and Titan, Saturn's largest moon, has been under study since 1977. The 1977 study' established the feus- ibility of such missions for the 1980s and early 1990s without the necessity f r Jupiter gravity assist. developed the mission strategies including space- craft and probe conceptual design. The 1980 study (called Cronos, the Greek god equivalent to Saturn) addressed details of the late 1980s mission opportunities, identified some 1990s Jupiter gravity assist trajectories, and began an inter- action with a study of missions to the planets beyond Saturn.5

The 197E2*3 and 1979 x studies further

Major participants in the Cronos study in 1980 included the Jet Propulsion Laboratory (overall mission and orbiter design), NASA Ames Research Center (Saturn and Titan probe desien), and the Japanese Institute of Space and Aeronautical Science (probe pre-entry science design). The Cronos study and the companion outer planets study were funded by the NASA Office of Space Sciences. Interfaces were maintained with synergistic activ- ities funded by the NASA Office of Applied Space Technology in spacecraft system desien, Titan probe technology assessment, aerocapture technology development, propulsion technology development (solar and nuclear electric), and spacecraft cub- system technology development. The study also interacted with Deep Space Network planning funded by the NASA Office of Space Tracking and Data Systems.

__ *This paper is based on work performed at the Jet Propulsion Laboratory, California Institute of Technology, under Contract NAS 7-100, sponsored by tho National Aeronautics and Space Administration.

**Task Leader, Member of the Technical Staff

This paper summarizes the work of a number of people and is basically a survey paper. Acknowl- edgments are included in the text. Unless other- wise noted the contributors are employees of the Jet Propulsion Laboratory. The 1979 and 1980 studies are more cam letely documented in the Cronos Final Report. 8

The 1980 Cronos study was characterized by a large number of mission options includiy varia- tions in interplanetary transfer mode, probe delivery strateey, satellite tour design, science payload, flight and information system configura- tion, and technobey readiness. The remainder af this paper describes the options and their impli- cations for the mission.

11. Satuïn System Science

A . Science Objectives

The recent passage of Voyager 1 through the Saturn system has focused the scientific objectives of the Saturn Orbiter with Probes Mission. However the basic objectives remain unchanged from those defined in 1977,1 i.e.:

1. To investigate the chemical composition and physical State of the atmosphere of Saturn, the atmosphere and surface of Titan, the Saturn ring system and the Saturn satellites and

2. To investigate the structure and physical dynamics of the Saturn magnetosphere.

Table 1 amplifies the objectives. The abjec- tives were adapted from the Galileo Jupiter mission by G. Orton and B. Tsurutani of JPL, with addi- tional objectives relating specifically to the Saturn ring system and to Titan as an object of inquiry in its own right. In addition to the over- all objectives, a rationale for pre-entry science for both the Saturn and Titan probes was developed by J. Cuzzi and T. Shimazaki of hes; K. Hirao, M. Shimizu, A. Nishida, and H. Hirosawa of the Japanese Institute of Space and Aeronautical Sciences (ISAS); and Orton and Tsurutani of JPL.

The pre-entry science region ranges from the Saturn magnetospheric boundary through the therma- spheres into the upper stratospheres of Saturn and Titan, to concentrations of about or pressures of about 10-4 mb. The region contains areas which will be inaccessible to the orhiter, e.g., zones of interaction between the magnetosphericlsolar wind environment and the planetary atmospheric environment, between the rings and the magnetosphere, and between satellites and the magnetosphere. The primary scientific objective for the pre-entry science is to under- stand the composition and structure of the upper atmosphere and ionosphere of both Saturn and Titan, and the controlling processes. Other objectives are to understand the ionosphere-magnetosphere coupling and the energy balance of this region in temi of sources, sinks, and injection mechanisms

1

Table 1 Mission objectives

General science objectives of a Saturn system mission are

To investigate the chemical composition and physical state of Saturn's atmosphere and Titan's atmosphere

To investigate the chemical composition and physical state of the Saturnian ring system

To investigate the chemical composition and physical state of the Saturnian satellites, with emphasis on Titan

To investigate the structure and physical dynamics of the Saturnian magnetosphere

In each of these three general areas of planetary, satellite, and magnetospheric investigations, the fallowing paragraphs list more specific science goals in approximately descending order of importance

Planet science objectives

1. Determine the chemical composition of the atmosphere

2 . Determine the structure of the atmosphere to a pressure depth of at least 20 bars

3. Determine the physical properties of atmospheric aerosols and their spatial distribution, both vertically and horizontally

4 . Determine the radiative energy balance in the atmosphere

5 . Investigate the circulation and dynamics of the atmosphere

6. Investigate the upper atmosphere and ionosphere, including neutral and charged particle currents

7. Investigate the occurrence of lightning in the atmosphere

Saturn ring system science objectives

1. Characterize the composition and physical state of ring particles

2. Characterize the total mass and radial distribution of material in the rings

3 . Search for undiscovered elements of the ring system

4 . Characterize the radial, azimuthal, and vertical (i.e., out-of-plane) dependence of particle number density, particle size distribution, composition, and physical orientation. This is to include azimuthal and vertical inhomogeneities, as well as asymmetries and vertical variations of the plane

5 . Study the outer rings and inner satellites for evidence of ongoing processes, such as accretion, erosion, o r gravitational maintenance of sharp boundaries

6. Investigate the interaction between the ring system and the radiation belt and the planetary magnetic field

1. Investigate the properties of any tenuous gases within the ring system

Titan science objectives

i . Determine the chemical and isotopic composition of the atmosphere, with some emphasis on the state of complex organic compounds

2. Determine the structure of the atmosphere down to the surface

3 . Characterize the physical state of the surface

4 . Determine the nature of the atmospheric aerosols and their spatial distribution, both vertically and horizontally

5. Determine the energy balance in the atmosphere and at the surface

2

Table 1 Mission objectives (continued)

6. Investigate the circulation and dynamics of the atmosphere

7. Investigate the surface composition

8.

9. Determine the gravitational fields and dynamic properties of the interior

Differentiate major surface compositional units

10. Determine the properties of an intrinsic magnetic field and radiation belt, if they exist

11. Investigate the occurrence of lightning in the atmosphere

12. Investigate the upper atmosphere and ionosphere, including neutral and charged particle currents, the nature and extent of surrounding neutral o r ion gas clouds, and interactions with the magnetosphere and with any particles in a thin, extended ring system

Satellite Science objectives

1 Characterize the morphology, geology, and physical state af satellite surfaces

2. Investigate the surface mineralogy of satellites and determine the distribution of compositional units

3 . Determine the gravitational fields, magnetic fields, and dynamic properties of the satellites

4 . Study satellite atmospheres and ionospheres, extended gas clouds arising from the Satellites, and interactions with the magnetosphere

5 . Search for new members of the satellite system

Magnetosphere science ob.jectives

1. Determine the configuration of the magnetosphere (out to 150 RS) and characterize its temporal variations

2. Characterize the energy spectra, composition, and angular distribution of energetic charged particles as a function of position and time

3 . Characterize the temperature, density, composition, and angular distribution of plasma as a function of position and time

Determine the plasma lieve modes existing within and around the magnetosphere, and the instabilities responsible for generating the emissions

4 .

5 . Investigate sutellite-magnetosphere interactions

6. Investigate ring-magnetosphere interactions

and to better characterize the satellite- and ring-magnetospheric interactions.

. B . Scientific Investigations

The science objectives will be pursued by both remote and in-situ experiments. The preferred strategy is for a Saturn probe, including pre-entry science, to make measurements in Saturn's atmos- phere at least to the 20-bar pressure region. (This is the region where a water cloud might be expected to exist.) be a surface lander which will make pre-entry, atmospheric, and surface measurements. The orbiter(s) will provide a data relay link for bath probes as well as making remote observations of the Saturn system.

The Titan probe is planned to

Because the Saturn orbiter mission is distant in time, and because the Voyager data has not yet been incorporated in the study, no "straman payload" was selected. Instead, a list of possible instruments was developed and related to the science objectives. Table 2 is a list of orbiter instruments, Table 3 af Saturn and Tiran probe instruments, and Table 4 of pre-entry science instruments. mented and revised when the Voyager science results are known, bnz they have senred to guide the 1980 mission design work.

These lists will undoubtedly be aug-

FOT both orbiter and probes, all science instruments flown on Galileo are prime payload candidates. However, the Galileo science payload is considered to be too mass constrained for really excellent science return to be expected. There-

3

Table 2 Saturn orbiter science: candidate experiments

Experiments requiring stable platform

Solid state imaging (narrow angle) Thermal IR microwave experiment

Solid state imaging (wide angle) Photopolarimeter

Near IR mapping spectrometer

High-resolution IR spectrometer UV imaging

Thermal IR experiment X-ray experiment

Microwave experiment Radar altimeter

UV spectrometer

Experiments not requiring stable scan platform or requiring spinning platform

Dust analysis experiment Planetary radio astronomy

Energetic particle detector

Magnetometer Actual aperture radar

Plasma analyzer Gravity wave detector

Plasma wave sensor Ultrastable oscillator

Orbiter neutral mass spectrometer

fore, additional experiments are planned to be flown on the Saturn mission. Additional experi- ments considered to be prime payload candidates ire the thermal infrared and the microwave experiments.

Experiments such as the radar altimeter and the actual aperture radar will probably receive addi- tional emphasis in future studies because of the Voyager-discovered opacity of Titan‘s atmosphere. The rationale for the selection of the Titan robe instruments has been previously do~umented.~,~

FOUT concepts were developed far the pre-entry science package by H. Hirosawa of ISAS. ISAS is considering providing pre-entry science packages for the probes. The most attractive option from their paint af view is a separate, completely integrated package of experiments which would relay its data either to the main probe or to the orbiter before being burned up in the atmosphere. The second and third concepts are for a package to be mounted on the probe (internally or externally) which would be jettisoned after its data-taking period to reduce the mass which the probe entry system would have to decelerate. The fourth option is to include the pre-entry science instru- ments as an integral part of the probe science. ISAS participation in the Saturn orbiter study is continuing in 1981.

111. Mission Profile

Many basic mission profile options were anal- yzed in the FY’77 and ‘78 studies. These and additional options were developed and comparative analysis was done in the N ’ 7 9 study, principally by Ronald Boain, James Gercchultz, and Phillip Roberts.

A. LaunchIInterplanetary Transfer Options

Figure 1 summarizes the interplanetary transfer options which have been considered for the Saturn orbiter mission. The space shuttle is the basic launch vehicle far all these options. studies of expendable launch vehicles have indi- cated that same missions are also feasible with the Titan-Centaur.) All the shuttle-launched Saturn missions require an additional two o r three stages for interplanetary transfer. Second-stage options include Centaur; the twin inertial upper stage (IUS) plus an additional, small IUS engine; or on- orbit assembly of several IUS engines. This latter option requires multiple shuttle launches and in- space assembly af from 3 to 6 large IUS engines plus the spacecraft and third propulsion stage (e.g. SEPS, if required). in 1980 has indicated that this concept is technically feasible.

(Limited

A small study9 conducted

Ballistic options to Saturn include a direct transfer, Jupiter gravity assist (JGA) i Venus-Earth gravity assist (VEGA); and a chemically powered Earth swingby called AV Earth gravity assist, o r AV-EGA. Sola r electric propulsion (SEP) can be used as a third stage and the options for it include a direct flight,solar electric Earth gravity assist (SEEGA), o r two Earth swingbys with SEP (called dual SEEGA or D-SEEGA). The solar electric options have been described in detail by Sauer. 10 only available in the mid-to-lute 1990s were studied by Uphoffl1 and compared with the other options. A few opportunities using JGA plus a AV-EGA were also investigated by Science Applications, Inc. and David Ross.

Jupiter gravity assist options which are

4

Table 3 Probe science and targets

Experiments

Gas chromatograph (S&T)a

Macs spectrometer (S&T)

VISIIR net flux and spectral radiometer (S&T)

VISIIR spectrometer (S&T)

Nephelometer and differential thermal analyzer (S&T)

Particle size spectrometer (S&T)

Atmospheric stnxtuïe experiment (S&T)

Energetic particlellighting detector (S&T)

Light element abundance experiment (S&T)

Impact accelerometer (T)

Radar altimeter (T)

Descent imaging (T)

Surface imaging (T)

Surface meteorology (T)

Surface sample acquisition and analysis (T)

a backscatter experiment (T)

Stable oscillator (S&T)

a s = Saturn: T = Titan

Nuclear electric propulsion (NEP) is another, langer term conce t which was investigated by an OAST-funded studye2 in 1980. It shows great pro- mise for delivering large payloads to Saturn with relatively short flight times, but will not be available until the 1990s.

Table 5 summarizes the launch energy require- ments far the iaunchltransfer options available in the 1980s. Comparison of the performance of the various launch/interplanetary options is made in a later section.

B. Orbit Insertion Options

Orbit insertion at Saturn can be done Conven- tionally, by firing chemical rockets, or by flying a lifting body through the atmosphere of Saturn or Titan to dissipate the interplanetary transfer energy as heat (aerocapture) ,13314 or by spiraling in using nuclear electric propulsion. Whereas solar electric propulsion systems are ineffective at the distance of Saturn from the Sun and cannot

Table 4 Saturn and Titan probe pre-entry science packages

Sat"=" (Titan) Instrument -

X

X

X

X

X

X

X

~~

(X) Ian mass spectrometer

(X) Neutral mass spectrometer

(X) Retarding potential analyzer

Electron temperature probe

Impedance probe (X)

(XI Magnetometer (search coil)

Energetic particle detector

Plasma analyzer

Plasma wave spectrometer

Table 5 Launch energy requirements f o r 1986-1989 launch

Trajectory type Launch C3 (km2/s2)

Direct ballistic 115 to 125

VEGA 15 to 23

AV-EGA 48 to 50

SEPS direct 40 to 60

SEEGA 10 to 12

D-SEEGA 1 to 2

be used for orbit insertion, nuclear electric pro- pulsion has no such restriction. Conventional orbit insertion can be performed using solid or liquid propellants. In 1980 the Cronoc study considered solids and two types of liquid rockets: "Earth-storable" monopropellant systems and "space-storable" hydrazine-fluorine systems. l5 Combinations of solids for large burns (the powered Earth swingby for AV-EGA and the orbit insertion) and liquids for trajectory correction maneuvers were also considered. Solid motors were assumed to have an ISp of 300 seconds, Earth storable systems 300 seconds, and space storable

5

( SEPS ) (BALLISTIC) (BALLISTIC)

AV-EGA JGA

Fig. 1 Interplanetary transfer options.

systems 370 seconds. with Earth storablec, or an all-liquid, space- storable system appear most attractive.

A combination of solid motors

The propellant mass requirements far the chemical orbit insertion increase exponentially with Saturn approach velocity. against high approach velocities and means in turn that long interplanetary flight times (6 to 8 years) ure required to approach Saturn slowly. On the other hand, the mass required of an aero- capture aeroshell increases approximately linearly with approach velocity, allowing shorter flight times for the same payloads for some trajectories. Flight times may be reduced to 5 years 01 less by using aerocapture for orbit insertion. tional advantage of aerocapture is that the inser- tion can be into a much lower period orbit than far a chemical system, reducing the time required to achieve scientifically desirable orbits.

This militates

An addi-

A 100-k1i nuclear electric propulsion system would allow flight times to Saturn as short as 4 years, including the time to spiral into a low periQd orbit. However, the detailed design of such missions was not pursued in the 1980 Cronos study.

C. Saturn Delivery Options

The Saturn orbiter mission can be flown with many different combinations af orbiters, probes, and Flybys. In addition, a number of strategies for deploying the probes are passible. Table 6 lists the various options which were considered in the FY'79 study. The Cronos study in FY'80 focused on SO2P (single orbiter carrying two probes) and SOSP/SOTP ( a split launch with two orbiters, each carrying a probe) in order to study orbiter design options. In order to allow compar- ison, delivery options studied used a Saturn approach speed of V, = 6.5 km/cec, an initial orbit period of 160 days, and u Saturn ring plane crossing constraint of 2.7 Saturn radii. (The aerocapture cases were exceptions.)

Far all options the orbiter or flyby bus was assumed to provide a data relay link for the probe. This constrains the flyby or orbit insertion tra- jectories. overflight is when both probes are deployed from the incoming asymptote and supported by a single orbiter or flyby bus. However, this is also the lowest mass case for SOZP o r S82P (a flyby bus carrying two probes). developed a strategy for this case. probe is deployed at 100-150 days out. The spacecraft is then targeted to Titan and the Titan probe is deployed 30-50 days out. Then the space- craft is retargeted to overfly the Saturn probe. The Titan probe enters first and is supported by the orbiter on its approach to Saturn. Saturn probe overflight is performed j u s t before Saturn orbit insertion.

The most difficult case for probe

In W'80 Chauncey Uphoff The Saturn

The

A navigation analysis performed by Jordan Ellis and Phil Laing evaluated requirements for the most difficult navigation cases: aerocapture in Titan's atmosphere and Titan probe deployment. ysis indicated that optical navigation is required for bath cases and that the Titan probe entry affer Saturn closest approach is extremely difficult. This latter circumstance is because trajectory errors are greatly magnified by Saturn's gravity at closest approach. General characteristics of the probe delivery and overflight trajectories have been documented previously. 1,2 trajectory was refined by William Blume for a FY'80 request for proposal developed by h e s for a Saturn Probe study contract. l6

The anal-

The Saturn probe

D. Saturn Tours

After orbit insertion the spacecraft will orbit Saturn for (nominally) 3 years. Two different types of "tours" were designed by Roger Diehl in FY'80 to al low an exhaustive study of Saturn and its environment and satellites. The tours were designed with science observation ground rules similar to those of the Galileo tour at Jupiter."

6

Table 6 Saturn delivery options

Option Description

1. S02P

2. SOTP

3 . SOSP

4 . so

5. S02P

6 . so2P

7. SOTP

8. SOTP

9. SOTP

10. SB2P

11. SBTP

12. SBSP

With SP deployed on approach and TP deployed from orbit

With TP deployed from orbit

With SP deployed on approach

With no probes

With both probes deployed on approach. SP enters 70"S, TP enters before Saturn periapsis (early)

With both probes deployed an approach. SP enters 3 - 5 , TP enters after Saturn periapsis (late)

TP deployed on approach and enters after periapsis (late)

Aerocapture with Titan V, = 7 h l s

Aerocapture with Titan V, = 10 kmls

Like option 6 using bus instead of orbiter

Titan entry far V, = 6 kmls

Like option 3

The tours were not designed to be optimal but were able to:

1. Show feasibility and evaluate AV requirements,

2 . Illustrate the differences between Jupiter and Saturn tours, and

3 . Identify possible pitfalls in accomplishing saturn t0"TC.

There are two major differences between Jupiter and Saturn tours, viz:

1. Four major satellites are available at Jupiter for gravity assist whereas only Titan is massive enough to rotate the orbiter's trajectory and change its inclination at Saturn.

2. The Jupiter radiation hazard is not present at Saturn but a periapsis constraint (nominally 3 Rs) is presented by Saturn's rings.

A "Titan-intensive tour'' incorporating targeted encounters only with Titan, and a "satellite-intensive tour" featuring close encounters with 5 af the mast interesting satellites were designed. These are illustrated in Figures 2 and 3 and their characteristics are summarized in Table 7. Table 8 illustrates the improvement in satellite observations which will be achieved over Voyager. The Titan-intensive tour maximizes the amount of rotation in Saturn's orbit plane (desirable for particles and fields science), while the satellite-intensive tour creates very close encounters with satellites other than Titan. A close encounter with Iapetus is very difficult because of its inclination (~15') to Saturn's equator. However, a zero AV Iapetus encounter opportunity was successfully identified in the

satellite-intensive tom. confidence in the feasibility of conducting Saturn tours for a reasonable AV budget (see the performance section, following).

The tour design gives

The major events in the two tours are orbit insertion, periapsis raise maneuvers, orbit period decrease, Titan probe deployment and overflight, and satellite encounters. In order to rotate the orbit and to change its period for other satellite encounters, Titan is encountered an nearly every orbit. This means that Titan can be almost as thoroughly mapped with radar as if a Titan orbiter were flown.

IV. Mission Options

The FY"1'9 study addressed a wide spectrum of orbiterlprabelflyby delivery options but analyzed only a limited number of flight system designs. The N'80 Cronos study focused on two basic delivery options, but analyzed and made cost ecti- mates for several designs. Consequently, the groundwork hac been laid for preliminary design of a Saturn orbiter flight system. study also derived reasonably solid performance requirements, the basic for proceeding into a project-level mission design is available.

Since the FY'80

The nominal launch date for the Saturn orbiter mission is currently 1989. Since JGA opportunities and nuclear electric propulsion technology will not be available until the mid-to-late 19YOs, the FY'80 study focused on the other trajectory options out- lined in an earlier section.

The design options selected for the FYI80 Cronos study are summarized in Table 9. orbiterlprobe options were stressed because they offer both the most opportunity for science return

The

7

1

NUMBER OF ORBITS

~ i ~ . 2 Titan-intensive tour 1980- '81. Fig. 3 Satellite-intensive tour,

Table 7 Titan-intensive vs satellite-intensive tours

Petal Inclined

Tour encounters satellites Magnetotail Dusk occultations (i > 2 0 " ) Number of Targeted Saturn ring orbits

Titan- 27 Titan Yes Yes Yes Yes

Satellite- 26 Titan Yes NO Yes Yes

in tens ive

intensive Dime Rhea Hyperion 1apetus.

Table 8 Satellite closest approach altitude (km) comparison

Voyager Titan-intensive Satellite-intensive Satellite f lybyc to"= tour

Mimas 88,796 43,805 57,752

Enceludus 86,569 48 ,248 21,566

Tethys 92,600 12,822 28,161

Dime 160,873 10,255 500

Rhea 71,997 45,514 531

Titan 4,109 337 258

Hyperion 470,714 35,375 1,001

1apetus 922,000 842,016 2,470

8

and the biggest design challenge. The target masses in Table 9 correspond to the mass which the target interplanetary trajectory mode can deliver to Saturn. The macs estimates and available tra- jectories show that, in most cases, the target masses are very difficult to achieve. Three major mass drivers were analyzed: the spacecraft mass, and the probe masses.

A. Performance Requirements and Characteristics

the AV requirements,

Propulsion mass is the largest component of the flight system mass. This results primarily from the mass required for orbit insertion. For the aerocapture mode the ueroshell is also the major mass component. Table 10 summarizes the AV requirements for four Saturn delivery options. These are exclusive of the SEPS o r the AV-EGA pro- pulsion and reflect the requirements after the SEPS o r the AV-EGA propulsion module has been released. The values far Saturn probe and Titan probe deploy- ment are far the deflection maneuvers to target the probes. The orbit insertion and Saturn tour values are based on the two tours designed in Fï '80. The navigation requirements include both deterministic maneuvers (calculated by Roger Diehl) and statistical maneuvers (calculated by Lanny Miller).

The orbit insertion AV is a function of the approach velocity and can be reduced by longer flight times. The figures in Table 10 are for V, = 6.5 kmlsec. Conversely, if mass margin is available because of launch vehicle performance

improvements o r spacecraft probe mass reduction, the flight time can be shortened by adding orbit insertion AV.

Table 11 summarizes the performance character- istics of the various interplanetary transfer op- tions for 1986-1989 launches. In general, perfor- mance degrades with time because of Saturn's increasing declination.

B. Spacecraft Options

Three basic spacecraft designs were studied in FY'79-'80. copy of the dual-spin Galileo spacecraft. That is, it has a spinning section containing particles,and fields instruments and a despun scan platform for the imaging science and other instruments requiring a stable platform. The "advanced" configuration is a 3-axis stabilized design utilizing technology which is anticipated to be available in 1985-'86. Douglas Turner led the benchmark and advanced spacecraft design efforts in FY'80. The "innova- tive" design is a spinning spacecraft with inter- nally spin-compensated optics which minimizes mass and uses technology which will probably require accelerated development to be ready for a 1989 launch. Roy Kakuda designed the innovative space- craft.

The "benchmark" configuration is a

Tables 12, 13, and 14 give mass estimates for the three spacecraft configurations. The science payloads were selected from the experiments listed

Table 9 E'Y'80 Cronos study - mission options

Achievable Configuration strategy trajectory Saturn approach, kg mass estimate, kg trajectory

Target Target mass FY'80 studya

Benchmark (GLL inheritance)

Benchnark

Advanced design (GLL 3-axis)

Advanced design

Innovative design

Innovative design

SOSPISOTP

S02P

SOSPISOTP

S02P

SOSP/SOTP

SOZP

( A ) AV-EGA 1,850 (8) SEEGA 2,500

Dual SEEGA 3,000

Direct ballistic 1,400 5 IUS stages

AV-EGA 1,850

Direct ballistic 1,000 STS Centaur

Direct ballistic 1,400 5 IUS stazes

2,110 SEEGA (6)b o r 3s-AV- EGA(7)

2,590 D-SEEGA (7)b or %-AV- EGA(8)

2,590 D-SEEGA (7)b or %-AV- EGA(8)

2,000 SEEGA (6)b or %-AV- EGA(7)

2,470 SEEGA (6)b or 3s-AV- EGA(7)

,780 AV-EGA (8) -

( ) Flight time in years

'At Saturn approach V_ = 6.5 kmlsec

bThese flight times assume V, > 6.5 kmlsec, i.e. that mass margin goes into orbit insertion AV.

9

Table 10 AV summary table ( m l s )

4.7 750

8 . 0 1450

7.9 1910

6 .8 2540

7.8 3860

Mission ootion SOZP SOSP SOTP so

4.7 650

8 . 0 1830

7 . 9 1860

6 .8 2520

7 .8 3830

~

Approach navigation 30 30 30 30 Reserve 10 10 10 10

sum 1 40 40 40 40 __ - -

SP deployment 25 25 - - Orbit insertion 1,150 1,150 820 820 Periapsis raise 230 230 275 275 Navigation 120 120 100 100 Reserve 30 30 25 25

Sum 2 1 ,555 1,555 1,220 1,220

TP deployment 100 - 100 - Satellite tour 100 100 100 100 Navigation 300 300 300 300 Reserve 50 5 0 5 0 5 0

Sum 3 550 450 550 450

Total 2,145 2 ,045 1,810 1,710

__

- -

4.7 1140

8 . 0 2150

7.9 2580

6 .8 4150

Table 11 Performance characteristics

4.7 1020

8 . 0 2730

7.9 2500

6 .8 4120

Launch year

Transfer Option

Direct

VEGA

IUS AV-EGA

SEEGA

O-SEEGA

Direct

30-k VEGA centaur

AV-EGA

SEEGA

%-Direct

3s-VEGA

35 AV-EGA

45 AV-EGA

On-orbit assembly

1986

Flight Approach time, mass, Y= kE

4.7 770

8 . 0 1890

7.9 2040

6 .8 2580

7.8 3920

4.7 1160

8 . 0 2790

7.9 2740

6 .8 4210

4.7 1620

8 . 0 3310

7.9 2910

7.9 4040

1987

Flight Approach time, mass, Y= kE

4.7 800

8 . 0 1250

7.9 1940

6 . 8 2560

7 .8 3890

4.7 1220

8 . 0 1870

7.9 2610

6 .8 4180

4.7 1710

8 .0 2200

7.9 2750

7.9 3840

time, mass, time , mass,

2730 7.9 2650

3800 7.9 3700

VEGA approach speeds are 5.7 , 5 . 7 , 5 . 6 , and 6.0 for 1986, ' 0 7 , ' 8 8 , and ' 3 9 , respectively, but performance at V_ = 6 . 5 kmls will be comparable.

Notes: 1. Approach velocity is 6.5 kmls except for VEGA transfer.

. 2. Interplanetary navigation AV, ACS gas, and spacecraft adapter masses are accounted for.

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Table 12 Benchmark spacecraft mass estimate (Galileo derivative; dual-spin)

Engineering subsystems, kg

structure 198.6

RF system 34.1

Madldemod 9.4

Power/pyro 141.8

Commldata 29.2

ACS 102.2

Cables 46.3

Temp 3 5 . 6

Devices 28.6

Data storage 8.9

sx antenna 5.6

Balluct/RRH 6.9

Assy hardware 3.9

Total 651.1

Science subsystems, kg

SSI/NA/i?A 42.3

NIMS 17.7

PPR 4.3

uvs 4.0

EPD 8.5

DDS 4.0

MAG 5.7

PLS 11.1

PWS 5.7

USO 2.0

SCAS 3.6

Total 108.9

Totals, kg

Engineering 651.1

8% 57.9

Science 108.9

3% 3.4

Probe adaptor 10.0

Total 831.3

Table 13 Advanced spacecraft mass estimate (1985-'86 technology; 3-axis)

Engineering subsystems, kg

structure 114. 8

RF system 32.0

Mod /d emod 7.9

Power/pyro 126.8

Coddata 29.1

ACS 50.0

Cables 44.3

Temp 35.6

Devices 31.3

Data storage 10.0

SX antenna 4.7

Assy hardware 3.9

Probe relay 6.0

Total 496.4

Science subsystems, kg

SSI/NA 28

SSIIiIA 25

NIEIS 18

PPR 5

T H I N 26

uvs 4

EPD 8

MAG 6

PLS 12

PIIS 4

uso 2

Total 138

11

Totals, kg

Engineering 496.4 20% 99.3 ___

595.7

Science 138.0 2 WL 27.6

165.6

Probe adaptor 10.0

Total dry mass 771.3

Table 14 Innovative spacecraft mass estimate (spin-stabilized spacecraft)

Subsystem Mass estimate range, kg

Communications 119 - 53

Tape recorder 19 - 19

structure 175 - 122

CDS 41 - 41

AACS 128 - 85

Motion compensation 34 - 34

Temperature control 17 - 17

Science 115 - 115

The mass estimates in Table 14 for the innova- tive design are based on projections of the feasi- bility af designing a spacecraft with such mass reducing features.

A concept related to the innovative configura- tion is the "dual mode" spacecraft. This concept uses the mass reducing techniques of integrated, body-fixed instruments, momentum wheels, and data storage. However, the spacecraft can be either spin-stabilized or 3-axis stabilized. During cruise and in times of particles and fields data taking the spacecraft would spin. During periods of target observations the spacecraft would be stabilized in 3 axes to point the body-fixed remote sensing instruments at the target. concept will require innovative sequence design to satisfy the requirements of both remote and païti- cles and fields instruments.

The dual-mode

Total 648 - 486

in Table 2. The benchmark payload is smaller than the advanced payload, even though the benchmark total mass is higher. This is because the dual spin configuration is less efficient in terms of the ratio of science mass to spacecraft mass. The benchmark payload is regarded as "adequate" by the study scientists and is very similar to the Galileo payload. The advanced payload is regarded as "goad.'

The advanced configuration is based on a 3-axis It includes a design done for Galileo in Fï'80.

spinning platform for particles and fields instru- ments. An alternative is for the instruments ïe- quiring spin to provide their own spinning plat- forms as part of the instrument. The 3-axis design is much less mature than the dual-spin design, and this is reflected in the differences in mass con- tingency s h o m in Tables 12 and 13. design incorporates an articulated scan platform for imaging.

The 3-axis

The innovative spacecraft design is much less mature than either of the other two. the goal af reducing mass to shorten flight time and/or enable more launch options, the design elim- inated as many mass drivers as passible. The major mass reducing features of the innovative design are:

In line with

1. No scan platform 01 articulated high-gain antenna. This means that real-time data trans- mission will not be available because the spacecraft must be pointed at the Earth to transmit. There- fore, a large amount of data starage is required.

2. Integrated science instruments. It is assumed, for example, that a single set of imaging optics can be designed to cover the range from the near IR through the Lyman-alpha region in W.

3. Internally spin-compensated optics. In order to achieve reasonable exposure times, partic- ularly in the IR, a counter-rotated mirror system must be used to keep a target in the field of view longer than the spin rate would allow for a body- fixed instrument.

C. Probes

The Saturn probe design used in the FY'79-'80

However, because of Saturn's studies was basically no different than the Galileo probe to Jupiter.18 lower gravity, the Saturn probe's entry velocity is much less than the Jupiter probe's, and, therefore, the heat shield mass is considerably less. A mass estimate of 270 kg was used for the Saturn probe with the science payload shown in Table 3. This is a somewhat optimistic number when a pre-entry science package (Table 4) is included. NASA h e s Research Center is in the process of procuring a design study16 for the Saturn probe, including accommodations far pre-entry science and the 20- bar design requirement. mass reduction.

The study will emphasize

A Titan probe design study was conducted in FY'79 for h e s by the Martin-Marietta Corporation.8 Three types of probes were designed: an utmosphere- only probe, a hard lander with surface sampling, and a soft lander with extensive surface sampling. The hard lander was used in the FY'79-'80 mission stud- ies and a mass of 228 kg was estimated, including a pre-entry science package. A technology assessment for the design was performed by Martin in FY'80 using NASA OAST funds.19 This assessment identi- fied four primary technologies needing additional development:

1. Surface sample acquisition system

2. Battery power system

3. Long life materials

4. Data storage devices (especially bubble memories)

The assessment also identified additional areas for systems study, i.e.,

1. Landing system stability and minimization af effects on surface conditions

2. Pre-entry science integration

3. Reliability 4 . Magnetically levitated momentum wlieels to

allow long missions with reasonable propellant loads. 4. Mass reduction

12

D. EiasslPerformunce Comparison

Table 15 summarizes the FY'80 mass estimates for four Saturn delivery configurations and three design options. The benchmark masses were reduced by about 1000 kg from the PY'79 to the FY'80 esti- mates because:

1. The FY'80 Saturn tour design allowed better, less conservative, propulsion estimates to be used, and

2. Space-storable propulsion (Isp = 370 sec- onds) was assumed in FY'80 in place of Ear th- storable (Isp = 300 seconds).

The advanced configurations show only modest mass improvements over the benchmarks, but carry a better science payload. The innovative configura- tions assume that significant mass reductions can also be made in both the Saturn and Titan probes (from 270 and 228 kg to 200 kg) . Both approach mass and mass in orbit are shown.

Figure 4 compares the in-orbit masses with the performance of a number of the launch-interplanetary transfer options. This figure was adapted from Uphoffll by William Blume. Uphoff analyzed Jupiter gravity assist (JGA) opportunities in the 1990s. The 1996 AV-EGA JGA opportunity was added by Blume as an example of a launch opportunity With higher payload capability. The aerocapture SEEGA case is shown because it features the best mass delivery performance. For Some other options the aero- capture performance in terms of short flight time/ high payload is no better than using chemical retro- propulsion. The advanced SEP case is based on FY'79 estimates and is now believed to be optimistic. On- orbit assembly cases and Centaur cases are not shown

SOSP

hark Advan Innov

831 771 500

216 209 177

796 745 515

270 270 200

- - -

2113 1995 1392 2886)

(95) (163) (130)

1185 1109 766

SOTP

Bmark Advan Innov

831 771 500

213 207 178

770 727 522

- - -

228 228 200

2042 1933 1400 (2983)

(90) (158) (130:

1215 1138 789

TOTAL TRIP TIME. ""

Fig. 4 Performance comparison.

required for a AV-EGA alone to be a viable option. For a split launch Table 9 shows that a 5-year, direct ballistic mission is feasible for the inno- vative design using on-orbit assembly of 5 IUS stages.

E. Design and Technalany 1 s s . u ~ ~

The major mission drivers on design and tech- nology are:

1. High science return desires

2. Severe mass limitations

3. Long mission duration (7 to 11 years)

4. Long communication distances (10 AU)

5. Cost limitations

Figure 4 and Table 9 illustrate that either an-orbit assembly or a SEPS is required by both tne benchmark and advanced configurations. The extremely low mass of the innovative design is

Table 15 Mass summary table (kg)

Mass element SOZP

bark Advan Innov

841 781 500

245 238 199

1005 950 677

270 270 200

228 228 200

2589 2467 1776 3550)

(127) (195) (160)

1264 1186 813

so

;mark Advan Innov

821 761 500

190 185 160

609 569 398

- - -

- - -

1620 1515 1058

Spacecraft

1nertsa

Propellanta

Saturn probe

Titan probe

Macs approach (FY'79 estimates)

Contingency

Mass in orbit

(58) (126) (100)

1144 1071 747

"Space-storable propulsion, I = 370 s SP

1 3

The design of the end-to-end information system (including ground, spacecraft, and probe information systems) presents particular difficulties. In FY'80 Kenneth Savary developed alternate end-to-end infor- mation system (EEIS) concepts for the Saturn orbiter mission. Because af the long flight time and commu- nication distances, the spacecraft must be more self maintaining than previous systems to have acceptable reliability and assurance of returning science data. As a compromise between cost and spacecraft mass limitations, the best telecommunication system de- sign appears to include a 67.2-kblc downlink and the use of an-board data compression, coding, processing, and storage. A large portion of the processing should probably occur in the instruments themselves. The EEIS design assumed that arraying of large Deep Space Network (DSN) antennas on the ground will be available to increase the downlink capability. The target for spacecraft data storage is 1010 bits which implies the use of advanced bubble memories or disk recorders. Variable rate data compression with Reed Solomon coding can be used to maximize science data return while protecting against errors. Elec- tronic beam steering andlor Ka band for high-rate downlink, and X-band uplink for high-rate commanding of the spacecraft appear to be desirable technolo- gies but require research and development.

During FY'80 telecommunication system analyses were conducted which investigated effects of antenna size and pointing accuracy, frequency, weather, and the interaction of telecom system design with data compression and coding. While Ka band or other very high frequency is desirable from a data rate point of view, X and S band are much better for radio science measurements in the Saturn and Titan atmospheres. X and S band are also much less affected by terrestrial weather than Ka band. In addition, electronic beam steering may be required for the high pointing accuracies required by Ka band. At the other end of the scale, B study con- ducted by Terry Grant of NASA h e s Research Center indicates that low frequency (-500 NHz-L band) is needed f o r communication between the orbiter and the Saturn probe to meet the 20-bar requirement. For the innovative and dual made spacecraft, a single high-gain antenna would probably be used to track the probes as well as to transmit to Earth. This implies that an integrated radio system with three or mote frequencies (S, X, and L band, fo r example) might be needed. The feasibility of this concept will be explored in future studies.

Use of variable rate data compression with Reed-Solomon coding was identified as i promising set of technologies to reduce telecom system size and power requirements. Application of this com- pression and coding to non-imaging data is an area requiring research.

Because of the long flight times, the majority of the costs of the Saturn orbiter mission may be f o r ground operations. Three means of reducing these costs were recommended:

1, An EEIS which serve5 several missions, which c m therefore split operations costs,

2. Automation of the spacecraft control activ- ities to reduce manpuwer, and

3. Maintenance of a multi-mission planetary data base to reduce costs of ground processing.

14

Plas5 reduction is a key mission need. In addi- tion to the mass reduction design features of the innovative spacecraft, the following technologies appear promising far lower mass spacecraft:

1. Low-density composite structures (potential 30-35% reduction of the advanced configuration structure mass).

2. More efficient power generation (potential 20-30% reduction in RTG mass).

3 . Space-storable propulsion systems (10-12% less propulsion system mass than Earth storablec).

4 . Sensor integration - optics and antennae (could eliminate 10-12 kg in the advanced payload).

Spacecraft lifetime can be increased by the ad- dition of redundant components. However, this ap- proach adds mass. Possible options for increasing lifetime without major mass increments include:

1. Fault toierancelautonomous spacecraft maintenance.

2. Use of fiber optics rotation sensors and magnetic bearings to eliminate conventional gyros and spin bearings with limited lifetimes.

3 .

An additional technology area in which develap-

Powered-dom mode operation during cruise.

ment may be necessary is cooling for science in- struments. In particular, IR instruments must be very cold to detect radiation from cold planets and satellites. Cooling devices which incorporate radiators can also be used as heat sinks for cool- ing the space-storable propulsion system. Active devices, i. e . , miniature Joule-Thompson refrigera- tors, are also under development and may have great mass advantages over passive devices.

Finally, B design issue was identified in FY'79 which impacts the feasibility of using a SEPS as a third stage for an orbiter carrying two probes. Because of the length and mass of this launch can- figuration, the stress on the chuttlelIUS interface during launch exceeds the allowable limits. This problem exists for other missions, notably the International Solar Polar Mission, and will pre- sumably be solved during normal development af the space transportation system (STS). For a Saturn orbiter carrying only one probe this problem does not exist.

The Saturn orbiter studies have not directly addressed the need for STS R&D. However, the devel- opment efforts on SEPS, N E P S , ~ ~ aerocapture,20 Centaur, and on-orbit assembly of IUSs are being monitored so that performance predictions can be as realistic as possible. that it is essential for STS improvements to be made for planetary exploration to continue.

F. Mission Cost Estimates

Beckman" has illustrated

Very preliminary cost estimates were made in FY'79 by Russell Nugorski for the benchmark con- figuration in a variety of delivery options. In FY'80 Daniel Spadoni of Science Applications, Inc., Schaumburg, Illinois, provided preliminary cost estimates for the benchmark and advanced configura- tions under various assumptions of the inheritance from previous spacecraft designs. Tables 1 6 and 17

m i

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J

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1 7