A Realistic Interstellar Explorer
Ralph L. McNutt, Jr. and the Realistic Interstellar Explorer Team
G. B. Andrews, R.E. Gold, A. G. Santo, R. S. Bokulic, B. G. Boone, D. R. Haley,
J. V. McAdams, M. E. Fraeman, B. D. Williams, M. P. Boyle, (JHU/APL)
D. Lester, R. Lyman, M. Ewing, R. Krishnan (ATK-Thiokol)
D. Read, L. Naes, (Lockheed-Martin ATC)
M. McPherson, R. Deters (Ball Aerospace)
The Johns Hopkins University Applied Physics Laboratory
Laurel, MD, U.S.A.
Abstract. From observations and theory we know that the unshocked solar wind extends at least 80 AU from the
Sun but likely no more than ~100 AU in the region from which the local interstellar wind blows. The much larger
region of the shocked solar wind and heliosheath extend out to at least several hundred AU, and fast neutrals from
charge-exchanged supersonic solar wind protons disturb the very local interstellar medium to ~500 AU or more. Thus
to really understand the interaction of the solar wind with the local external medium, a properly-instrumented, in situ
probe to this region of space is required. For more than 20 years, an “Interstellar Precursor Mission” has been
discussed as a high priority for multiple scientific objectives. The chief difficulty with actually carrying out such a
mission is the need for reaching significant penetration into the interstellar medium (~1000 Astronomical Units (AU))
within the working lifetime of the initiators (<50 years). We have revisited an old idea for implementing such a
mission. The probe and its perihelion carrier are launched initially to Jupiter as a combined package and then fall to
the Sun where a large propulsive V maneuver propels the package on a high-energy, ballistic escape trajectory from
the solar system. Outbound in deep space, the two separate, and the probe takes data with its onboard instruments and
autonomously downlinks the data to Earth at regular intervals. The implementation requires a low-mass, highlyintegrated spacecraft to make use of available expendable launch vehicles. We provide a first-order cut at many of the
engineering realities associated with such a mission. These separate into (1) the systems constraints imposed on the
perihelion package by the combination of the propulsion system, carrying the needed propellant into perihelion, and
the associated thermal and mechanical constraints, and (2) the requirements of power, autonomous operations, and
data downlink from the probe itself. We find that many of the requirements for a low-mass probe that operates
autonomously for this mission are common for either this propulsion concept or more advanced low-thrust concepts,
e.g., solar sails and ion propulsion. We describe an implementation that could make such a mission into reality in the
next 10 to 20 years.
INTRODUCTION
MISSION CONCEPT
A mission past the boundary of the heliosphere has
been discussed for more than 25 years and would yield
a rich scientific harvest1-4. To the best of our current
knowledge, the external ionized interstellar wind flow
is supersonic with respect to the Sun. Hence, the
heliosphere will set up a shock wave in the local
interstellar medium, and the external shock may be as
much as ~300 AU away at its closest. So 1000 AU is
“clear” of the influence of the Sun on its surroundings
and is a reasonable distance goal for a probe mission
to the very local interstellar medium5-10.
The goal for the mission is to reach this distance
within the working lifetime of the probe developers
(<50 years) using a solar gravity assist (due to Oberth,
1929)11. In addition, launching toward a star enables
comparison of local properties of the interstellar
medium with integrated properties determined by
detailed measurements of the target-star spectrum.
Salient mission features include: (1) launch to Jupiter
and use a retrograde trajectory to eliminate angular
CP679, Solar Wind Ten: Proceedings of the Tenth International Solar Wind Conference,
edited by M. Velli, R. Bruno, and F. Malara
© 2003 American Institute of Physics 0-7354-0148-9/03/$20.00
830
operational capability. The latter implies open-loop
control, software autonomy, and autonomous safing
and recovery. These features could allow possible
extension to multi-century flight times while
maintaining data taking and downlink operations.
Appropriate use of redundancies could extend probe
lifetime to >1000 years (~20,000 AU). The model here
is the nominal five-year design life of Voyager 1 and 2
that has now been exceeded by a factor of five.
momentum, (2) fall into 4 solar radii from the center of
the Sun at perihelion, and (3) use an advancedpropulsion system ∆V maneuver to increase probe
energy6,7,9,10. The star ε Eridani was selected as the
target (Fig. 1).
To reach ~20 AU yr–1, the probe needs to be
accelerated by ~10 to 15 km/s during about 15 minutes
around perihelion6,11. Such acceleration was provided
to Ulysses in Earth orbit with a 20-metric-ton, 3-stage
solid-fuel rocket. To accomplish the same thing near
perihelion a much lighter, i.e., high specific impulse,
system is required.
We considered several propulsion options5-7,9,10.
Nuclear pulse propulsion (“Orion”) offers very good
performance but does not scale to small systems.
Nuclear thermal propulsion (NTP) for small systems is
limited to solid-core reactors and the Isp is limited by
fuel pellet coefficient of thermal expansion (CTE) and
chemical reactivity. Solar thermal propulsion (STP) is
being investigated for use in orbital transfer vehicles at
1 AU, although use at 1 AU requires concentrators.
For this mission, the location of the “burn” eliminates
the need for concentrators (albeit replacing it with a
need for a very-high-temperature thermal shield/heat
exchanger) while potentially avoiding the inherent
CTE problems in a solid-core NTP system. Hence,
STP is the propulsion option we have implemented 12.
FIGURE 1. Trajectory and mission event-dates.
ENABLING TECHNOLOGIES
The enabling technologies
link science,
instruments, the spacecraft engineering, and
reality6,7,9,10. To keep the mission based in a realizable
infrastructure, we use Evolved Expendable Launch
Vehicle (EELV) capabilities. Maximum capability is
provided by the Delta IV 4050H + Star 48B that we
have selected.
Trade Studies
We have concentrated on an STP system12, but
most results also apply to NTP. To obtain a
sufficiently large Isp to provide ∆V, we have examined
LH2, CH4, and NH3 and maximized the propellant
temperature (up to structural failure). Trades include
pressure versus flow rate, heating, and recombination.
LH2 was selected over NH3 as primary STP system
propellant. The higher NH3 density allows for smaller
thermal shield mass, but the higher Isp for LH2 allows
more packaging room for probe and room for
“growth” with lighter materials. The requirements for
heat exchanger coating are driven due to erosion of
carbon-carbon by all of the studied propellants. We
sized the propellant tank for propellant requirements,
including considerations of storage for cruise and the
need for a cryostat required for long-term LH2 storage,
as well as requirements on pressure and expulsion
during the propulsive maneuver.
Innovations that are needed include: high Isp, highthrust propulsion (for perihelion maneuver, ~15
minutes), a carbon-carbon thermal shield, a technology
that can be proven with the Solar Probe mission, and
long-range, low-mass telecommunications. The
extreme distance and low available power requires an
optical downlink.
For power we need an efficient Radioisotope
Thermoelectric Generator (RTG). For stable operation
and a long extended mission life we can use proven,
safe technology with Pu-238 plus extended operations
with Am-241 in the same RTG container.
To enable the fifty-year design lifetime and keep
options open for significantly longer operations, we
need low-temperature (<150K) electronics that
inherently provide long life with minimal
heaters/insulation as well as fully autonomous
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using ultra-low power (ULP) electronics operating at
cryogenic temperatures (Table 1).
Spacecraft Systems
The spacecraft consists of a pre-perihelion carrier
that is dominated by the thermal shield, cryostat, STP
system, and the instrumented probe (Fig.2). Following
the perihelion “burn” the carrier is discarded and the
instrumented probe deploys instruments and the main
communications optic (Fig. 3), verifies its status,
established the optical communications downlink13,
and begins its long cruise/data-taking mission.
TABLE 1. Model Science Instrument Payload.
Instrument
Mass (kg)
Power (W)
Totals
12.16
1.87
Magnetometer
1.89
0.18
Plasma wave
1.48
0.08
Plasma spectrometer
0.97
0.53
Lyman alpha imager
3.43
0.13
Cosmic ray spectrometer
0.84
0.16
Energetic particles
0.80
0.33
Dust experiment
0.70
0.13
X-ray spectrometer
2.05
0.33
Hydrazine G&C tanks (2)
LH2 cryostat
The mass and power of the probe have been
estimated in some detail. The mass estimate by
subsystem is 147.15 kg using a Be structure. The
power estimate by subsystem is 19.81 W using ULP
(Table 2). Special consideration has been given to
lightweighting the structure and all other components,
minimizing the power, and maximizing the probe
lifetime such that the mission requirements can be met
with the specified EELV.
Tertiary shield
Primary
shield/STP
exchanger
Secondary
shield/STP
exchanger
Stowed Probe
(within adapter)
2D STP
propulsion nozzle
FIGURE 2. Probe and thermal-shield carrier in preperihelion configuration.
For example, with respect to avionics we eliminate
signal harness and maximize autonomy by using
multiple ULP processors, each configured with two
wireless communication transceivers.
Deployed Optic
Attitude
Control
GN2 (3)
Stowed Optic
X-ray photometers
We use RTGs based upon multicouples with four
General Purpose Heat Sources (GPHS) and direct
voltage (no DC/DC converters). Am-241 is identified
as available RTG fuel: 1 Pu-238 RTG + 2 Am-241
RTGs provide for 51.3 W remaining after 50 years and
~6.27W after 500 years (all three RTGs).
Low-gain RF
antenna
Plasma
Wave antenna
deployer (2)
Ly α imager
An integrated optical communications, attitude,
guidance, and control approach ensures required burstmode performance of 500 bps from 1000 AU (over 5
light-days). Already at 100 AU, electromagnetic
waves take 13.9 hours to travel from the spacecraft to
Earth, so full open-loop autonomy is required.13.
RTGs (3)
FIGURE 3. Probe configuration during cruise
The mass estimate of the payload is 12.16 kg using
Be structure. The power estimate of payload is 1.87 W
TABLE 2. Probe Mass and Power
Component
Mass (kg)
Science instruments
Integrated avionics
Power system
Telecommunications
Structure
Attitude control
Thermal
Propulsion
Dry mass/power total
Total with harness
Power reserve
Total for probe including reserve
12.16
1.44
45.60
16.17
18.32
5.40
0.00
41.49
140.58
147.15
0.00
147.15
Power
(W)
1.87
1.38
2.12
11.28
0.00
1.69
0.00
0.16
18.50
18.87
0.94
19.81
832
Notes
Allocated total is 10.0 kg; power is nominal, not peak
Includes command and data handling, guidance and control
Three RTGs: 1 with Pu-238, 2 with Am-241
890 nm laser downlink plus RF for inner solar system
Beryllium truss
Attitude fine control integrated with laser downlink
All thermal input via passive waste heat from RTGs
Cold-gas N2 system warmed by RTG waste heat; 56W max
Subtotals of components before harness
Signal harness is virtual with RF; power harness use wires
Add 5% power reserve
Probe totals including reserves
Thermal requirements include: survive cruise prior
to perihelion pass, protect the propellant, heat the
propellant for the perihelion “burn, ” and use waste
heat from the RTG to eliminate heater-power.
Analysis with Thermal Synthesis System (TSS)
software shows the carbon-carbon (CC) primary shield
can be limited to a temperature (2924K) if we use
~100 kg of CC aerogel backing on shield while the
LH2 remains thermally isolated until needed.
ACKNOWLEDGMENTS
This work was supported under Task 7600-039
from the NASA Institute for Advanced Concepts
(NIAC) under NASA Contract NAS5-98051.
REFERENCES
1. Jaffe, L. D., and Ivie, C. V., Icarus, 39, 486-494, (1979).
2. Holzer, T. E., et al. The Interstellar Probe: Scientific
objectives for a Frontier mission to the heliospheric
boundary and interstellar space, NASA Pub., 1990.
3. Mewaldt, R. A., Kangas, J., Kerridge, S. J., and
Neugebauer, M., Acta Astron., 35, Suppl., 267-276
(1995).
FIGURE 4. Peak temperatures (degC) at perihelion.
4. McNutt, R. L., Jr., Gold, R. E., Roelof, E. C., Zanetti, L.
J., Reynolds, E. L., Farquhar, R. W., Gurnett, D. A., and
Kurth, W. S., J. Brit. Int. Soc., 50, 463-474 (1997).
SCHEDULE
5. McNutt, R. L., Jr., “A realistic interstellar explorer”,
Proc. Workshop on Interstellar Exploration in the Next
Century, 1998, California Institute of Technology.
Table 3 gives a schedule for implementation.
Date(s)
2000-2002
2002-2003
2003-2007
2004-2007
2004-2007
2007-2010
2009-2012
2012
2012-2015
2015-2065
35850
TABLE 3. Schedule.
Activity
Advanced technology development studies
Update NASA strategic plan
Development of small probe technologies
Development of solar-sail demo mission
Development of Solar Probe mission
Focused development of Interstellar Probe
Design and launch of solar-sail probe
Test Solar Probe performance at perihelion
Design and launch perihelion STP probe
Data return from out to 1000 AU
Probe reaches vicinity of ε Eridani
6. McNutt, R.L., A Realistic Interstellar Explorer, Phase I
Final Report, NIAC CP98-01, May 1999.
7. McNutt, R. L., Jr., Andrews, G. B., McAdams, J., Gold,
R. E., Santo, A., Oursler, D., Heeres, K., Fraeman, M., B.
Williams, B., “A realistic interstellar explorer”, STAIF2000 Proc. (2000).
8. Liewer, P. C., Mewaldt, R. A.,. Ayon, J. A., and Wallace,
R. A., “NASA's interstellar probe mission”, STAIF-2000
Proc. (2000).
9. McNutt, R. L. Jr., et al., “Low-Cost Interstellar Probe”
Paper IAA-L-0608, Fourth IAA International Conference
on Low-Cost Planetary Missions, Laurel, MD, May
2000; Acta Astronautica, in press, 2002.
10. McNutt, R. L., Jr., et al., “A Realistic Interstellar Probe”
COSPAR Colloquium on The Outer Heliosphere: New
Frontiers, Potsdam, Germany, July 24–28, 2000;
COSPAR Colloquia Series, 11, 431-434 (2001).
SUMMARY AND CONCLUSIONS
The initial concept5-7 for a robust, long-lived 50-kg
probe has proven difficult to design. A robust probe is
more likely to have a mass of ~150 kg, even with new
technologies and materials. Novel propulsion is key.
The continuing problem is the mass of the cryostat.
Due to the system dry mass, a 20 AU/yr goal may not
be realizable, although 12 AU/yr (60 km/s) may be.
More technical definition and engineering work is
required, but there is a road to actually implementing a
mission that has remained elusive for at least 25 years.
Ad Astra!
11. McAdams, J. V., and McNutt, R. L., Jr., “Ballistic
Jupiter gravity-assist, perihelion-∆V trajectories for a
realistic interstellar explorer,” Paper AAS 02-158, 2002.
12. Lyman, R. W., et al., “Solar Thermal Propulsion for an
Interstellar Probe” 37th Joint Propulsion Conference,
July 8-11, 2001, Salt Lake City, Utah.
13. Boone, B. G., R.S. Bokulic, R. S., Andrews, G. B., and
McNutt, R. L., Jr., SPIE Paper 4821-26, in press, 2000.
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