NASA Technical Reports Server (NTRS) 20050245135: Modular, Reconfigurable, High-Energy Systems Stepping Stones

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by  NASA Technical Reports Server (NTRS)

Modular. Reconfigurable, High-Energy Systems Stepping Stones
rAC-OS D3.2.04
56th International A~ronautical Congress
October 17-21, 2005
Fukuoka, Japan
Joe T. Howell
ASA Marshall Space Flight Center
Advanced Projects Team, SP20
Huntsville, AL 35812
E-mail: [email protected]
Tel: [phone]
John C. Mankins
Artemis Innovation Management Solutions, LLC
Ashburn,VA20147 USA
E-mail: [email protected]
Tel: [phone]
Connie Carrington
NASA Marshall Space Flight Center
Advanced Projects Team, SP20
Huntsville, AL 35812
E-mail: [email protected]
Tel: [phone]
Abstract
Modular, Reconfigurable, High-Energy (MRHE) Systems are stepping stones to provide
capabilities for energy-rich infrastructure strategically located in space to support a variety of
exploration scenarios. Abundant renewable energy at lunar or Eruth-Moon Libration (e.g., "Ll")
locations could SUppOlt propellant production and storage in refueling scenarios that enable
affordable exploration. Renewable energy platforms in geostationary Earth orbit (GEO) can
collect and transmit power to satellites, or to Earth-surface locations. Energy-rich space
technologies also enable the use of electric-powered propulsion systems that could efficiently
deliver cargo and exploration facilities to remote locations. A first step to an energy-rich space
infrastructure is a IOO-kWe class solar-powered platform in Earth orbit. The platform would
utilize advanced technologies in solar power collection and generation, power management and
distribution, thermal management, and electric propulsion. It would also provide a power-rich
free-flying platform to demonstrate in space a portfolio of technology flight experiments.
This paper discusses the reasons why such advances are important to future affordable and
sustainable operations in space. It also presents preliminary concepts for a IOO-kWe solar
powered satellite with the capability to flight-demonstrate a variety of payload experiments and to
utilize electric propulsion. State-of-the-art solar concentrators, highly efficient multi-junction
solar cells, integrated thermal management on the arrays, and innovative deployable structure
design and packaging make the lOO-kW satellite feasible for launch on one existing launch
vehicle. Higher voltage arrays and power management and distribution (PMAD) systems reduce
or eliminate the need for massive power converters, and could enable direct-drive of high-voltage
solar electric thrusters.
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An Affordability Challenge typical EEL V (evolved expendable launch
vehicles) costs per launch at the same rate.
One of the central barriers to more • [n Space Transport: Assuming expendable
ambitious-yet still affordable-space systems involving at least two in-space
operations in the Earth's neighborhood lies stages with individual mass of about 10,000
in our inability to affordably preposition kg (and a recurring unit per kilogram cost of
consurnables (particularly propellants) and about $50,000 per kilogram ') for a "per
needed systems (including spares for in mission cost" of about $500M per mission.
space servicing and maintenance). As long • Excursion Transpoli: Assuming expendable
as it is not possible to locally repair and systems, involving nominally a descent
refuel high-value (and high-cost) space module and an ascent module with a
systems beyond low Earth orbit (LEO). This combined mass of about 10,000 kg to
challenge affects planning for a wide range 20,000 kg (and a recurring unit per kilogram
of potential future missions, but is cost of approxi mately $50,000 per
kilogram), for a "per mission cost" of about
particularly important for (a) major, high
$750M per mission.
value missions such as human exploration
activities beyond low Earth orbit; (b) large Transportation Operations. Assuming
scale defense and/or security focused incremental improvements on Space Shuttle
mission systems; or, (c) 'future space and International Space Station (ISS) era
ground operations concepts" involving
industries' (such as larger, multi-payload
nominally about 20,000 total personnel
geostationary Eroih orbit (GEO) platforms,
(with an average cost of about $) 00,000 per
space solar power systems, and related
FTE2 (full time equivalent)), for a "program
concepts).
per year cost" of about $2,000M, and a per
mission cost of about $500M per mission at
For example, a large-scale, permanently a rate of 4 missions per year, or about to
inhabited lunar base might involve 3-4 $667M per mission at a rate of3 missions
human missions to the Moon per year (for per year.
crew rotation every 120 days or 90 days,
respectively). However, if such a mission
scenario were to involve Apollo-era
concepts and current-technology expendable I The very rough, mass-based cost estimation
space transportation systems, then the total relationships (CERs) used in this illu tration are
intended to be consistent with-and perhaps a bit on
cost per mission due to transportation costs
the optimistic side of- past human-rated space
alone (hardware and operations), could
systems recurring hardware costs. A more detailed
range from $2,400M per mission to more analysis would consider the specific cost per
than $3,100M per mission (current year kilogram for each of the major elements in each
dollars). Transportation cost components sy tem, as well as taking into account the specific
number of unique elements to be manufactured over
here are assumed to include the following:
the life of the program, and the degree to which
• ETa Transport: Assuming Shuttle-derived, advanced production methods (such lean
manufacturing) might be brought to bear on the
expendable systems involving 2 Heavy Lift
problem. However, for the purposes of this
Launch Vehicles (HLLVs), 1 Crew
discussion, the single value of 'about $50,000 per
Launcher at a total cost of $500M to
kilogram (about $22,000 per pound) for recurring
$l,OOOM per mission. Note that this rough
flight hardware costs eems adequate.
estimate for ETO costs is intended to be
2 This figure is intended to be a very rough average,
comparable to (but lower than) the Space
integrating all personnel involved; clearly some cost
Shuttle at about 3-4 launches per year, plus
categories, sllch as senior engineers, are at mllch
higher labor rates when 'fully wrapped'.
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In summary, this scenario would result in an onboard autonomy. Four functional
annual cost-for IWlar base transportation challenges must be resolved to enable this
only-of approximately $7,000MIyear (best highly desirable, but technical difficult
case,3 missions/year), to about transition:
$ll,OOOMIyear (worst case, 4 missions per
• Lower cost ETO transport (perhaps by
year). Additional costs would, of course be
enabling a transition to launchers that are
incurred for crew transportation systems,
more similar to those used by other
supporting infrastructures (such as government organizations or by commercial
communications systems), as well as for the sectors; and in the long term by transitioning
wide range of surface systems that would be to reusable launch vehicles);
needed for a lunar base. (It is perhaps worth
• Highly-autonomous assembly, maintenance
noting that in the case of the tightly
and servicing of modular systems in space
interwoven Space Shuttle and ISS programs,
and on planetary surfaces (including both
the costs of transportation to and from the robotic and crew-assisted operations),
Station are very roughly equivalent to the
• Affordable and timely pre-positioning of
costs of ISS engineering and operations. If
fuel, systems and other materiel throughout
the same were to hold true for a far-more
the Earth-Moon system (including to the
technically challenging lunar base and its
surface of the Moon); and,
transportation system, then the total annual
• Reusable, highly reliable and high-energy in
costs would be double the figures noted
space transportation (and for lunar missions,
above-or equivalent something greater
excursion transportation systems).
than the entire current U.S. annual civil
space budget.) The total of such annual
The systems that would enable such
operational costs would, of course, far
visionary capabilities must also be highly
exceed the current annual US investment in
autonomous (to reduce ground operations
human space flight. Moreover, they would
costs), as well as substantially less
not allow for vitally needed investments in
expensive to buy and own (with greater
the systems that would allow us to go
operational margins than current systems, as
beyond and initial operational base.
well as lower per wlit costs-perhaps
acrueved through modularity and the
Although the sketch above is specific to a
economies of production).
notional lunar base, the affordability issues
involved are quite similar for a range of
Detailed studies would be needed to
other ambitious future space operational
determine the appropriate technical
scenarios- particularly those involving (as
performance objectives for such advanced
does the Terrestrial Planet Imager (TPI)
systems-in the context of cost constraints
concept) a number of exceptionally large,
and reliability (safety) goals. However, it
rugh-value imaging systems deployed
seems plausible to suggest that at a
beyond low Earth orbit (LEO).
minimwn, such future R&D efforts should
target new systems approaches and novel
A Notional Solution technologies that would make possible not
less that a factor of four reduction in per
One potential solution to this challenge is to
mission costs, and perhaps as much as a 10-
move successfully to more affordable
fold reduction. In the case of the lunar base
reusable space transportation system
example sketched above, a 4-fold reduction
elements with substantially higher levels of
would be equivalent to seeking to achieve a
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lunar base per mission transportation cost of technologies should be broadly beneficial to
no more than $1,750M to $2,750M per the full range of ambitious mission options
mission--or, in the case of a lO-fold that are under consideration by various
reduction, a per mission transportation cost organizations.
of no more than $700M to $l,lOOM per
mission. (For comparison, note that the Key Technical Challenges
latter figures are roughly comparable to the
fully-loaded costs of Space Shuttle missions The central functional issues associated with
to LEO at the present time- although they affordably realizing advanced, highly
are still much greater than the marginal costs reusable architectural concepts lies in
of such flights.) solving several key technology challenges.
These include:
However, setting a goal is hardly the same
• Tele-supervised (and eventually
as achieving it. Although the technologies
autonomous) highly resilient deep space
needed to achieve this vision are (in many systems operations (in this case, 'deep
cases) already validated in the laboratory, space' operations includes all ambitious
they are certainly not 'in hand' or mission operations beyond LEO).
sufficiently mature to incorporate into space
• Reconfigurable and self-adaptive modular
systems being build today. As a result,
systems.
substantial research and technology
• Space assembly, maintenance and
development and validation must still be
servicing (from the systems level, down
undertaken in order to realize the potential
to the subsystem level).
cost savings that are so clearly needed in
end-to-end space transportation. • Highly fuel-efficient, high reliability, re
startable propulsion, such as high-power
Fortunately, the capabilities for local electric propulsion for cargo and
refueling, as well as locally autonomous cryogenic engines for time critical
assembly, repair and maintenance are mission (such as those involving
inherent for any kind of extended and astronaut crews).
ambitious deep space scenario- such as a
• High-energy propellants for long
lunar surface base.3 Moreover, they become
duration missions (particularly cryogenic
even more critical as one considers the long
propellants such as liquid oxygen, liquid
term requirements of human mission to
hydrogen, etc.)
Mars, much less the far more ambitious
requirements of extended human presence • Long-term storage and management, as
and activity in space (e.g., space settlements well as the highly reliable and low-loss
or missions beyond the llmer solar system). transfer (including transfer in micro
As a result, the future development of such gravity) of cryogenic propellants.
• High-power, but low-mass space power
generation and management systems
3 Note: although smaller missions-such as those
involving traditional communications satellite -
might benefit from in- pace refueling, low-cost in
space transportation and similar space operations, the This paper will deal with a specific class of
costs of developing and deploying sllch these technology problems: in particular,
transformational new capabilities are difficult, if not those that involve affordable, high
impo sible to justify based on these missions alone.
efficiency in-space transport of logistics
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using modular, reconfigurable high-energy attachment at the central of gravity of the
systems. integrated vehicle.
A Novel Concept: MRHE However, at this time MRHE systems entail
considerably greater system development
Modular, reconfigurable high-energy uncertainty than more conventional systems
(MRHE) systems are one novel conceptual and technologies (e.g., fully expendable,
approach with the potential to meet the Apollo-era concepts with technologies that
challenge of enabling affordable pre are already at a flight-like technology
positioning of key logistics (including fuel, readiness level- i.e., TRL 7-9). As a result,
hardware, and appropriate systems) to points significant R&D investments are needed
beyond LEO. An MRHE system is an prior to beginning major systems
integrated, high power, solar electric development. However, if affordability and
propulsion (SEP) vehicle that provides high sustainability are important characteristics
fuel efficiency from LEO to destinations for future transformational space operations,
throughout the Earth-Moon system then the development of new technologies
(including MEO, GEO, low lunar orbit and capability is essential. MRHE systems
(LLO) and Sun-Earth Lagrange points). (or equivalent alternate approaches to
solving these strategic challenges) is one
The MRHE concept consists of identical possible approach.
solar-powered modules, each equipped with
an electric propulsion system, assembled in Initial MHRE Efforts
a reconfigurable arrangement that supports
power generation redundancy and engine A variety of technologies and concepts must
out capabilities. For example, a promising be developed to enable a lOO-kWe solar
MRHE approach is a linear configuration powered satellite with the capability to
such. as that illustrated in Figure 1 (shown in flight-demonstrate a variety of payload
a gravity gradient stabilized mode). experiments and to utilize electric
Alternatively, a tetrahedral configuration has propUlsion. State-of-the-art solar
also been examined, such as that illustrated concentrators, highly efficient multi-junction
in Figure 2. solar cells, integrated thermal management
on the arrays, and innovative deployable
Figure 3 illustrates respectively the full structure design and packaging make the
system and a in a normal-to-the-orbit-plane 1O O-k W satellite feasible for launch on one
configuration. And Figure 4 illustrates a existing launch vehicle. Higher voltage
typical self-assembly sequence for an arrays and power management and
MRHE. distribution (PMAD) systems reduce or
eliminate the need for massive power
In any case, the MRHE concept provides for converters, and could enable direct-drive of
payload attachments on each module to high-voltage solar electric thrusters.
provide flexibility and re-configurability
options for accommodating multiple An objective of current R&D efforts, led by
teclmology experiments and eventually the NASA Marshall Space Flight Center,
different exploration payloads. In the case with the strong participation of several other
of a linear configuration, the platform may organizations (including ASA centers
also provide a single, larger payload such as JPL and GRC- and industry- such
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as ENTECH and LM) is the development of that address the need for affordable, routine
a cluster of four "mini-sats" in order to exploration, deli vering cost savings through
address all functional and operational issues. multiple-unit production, and providing
The proposed concept consists of identical options for affordable module replacement
solar-powered modules, each equipped with and module redundancy. On-orbit assembly
an electric propulsion system, assembled in will provide mission planners more
a reconfigurable arrangement that supports flexibility in choosing launch vehicles and
power generation redundancy and engine support hardware, increasing mission value
out capabilities. The concept also provides and affordability. Figure 5 illustrates two
for payload attachments on each module to alternate approaches that have been
provide flexibility and re-configurability examined during the past decade (a
options for accommodating multiple monolithic square array and a monolithic
technology experiments and eventually 'bat-wing' array). The central question of
different exploration payloads. economic feasibility is whether a modular
approach can yield a substantially less
The goal of the project is to assess and expensive operational system.
mature the technologies that support future
development of a modular architectures and As illustrated in Figure 6, there is a great
systems; in the case of this project, a 100 potential for these high-power, advanced
kilowatt-class spacecraft suitable for on transportation systems to provide reasonably
orbit assembly and reconfiguration. The fast trip times for transport within the Earth
thrust of the technology maturation effort is Moon system. (For example, this involves
to integrate a representative advanced solar less than 4 months to transport payloads
power generation (SPG) system, high LEO to GEO, with additional time for return
voltage power delivery system (PDS), and of the transport system to LEO for refueling
advanced thermal management into system and reuse.) At the heart of any decision
of-systems demonstrations. The activity also whether or not to develop these systems is
will address multi -satellite rendezvous and the as yet unresolved question of what
self-assembly, as well as distributed wireless 'learning curve' to use in analyzing the
avionics and robotic reconfiguration in the expected cost of such systems in operational
event of a simulated spacecraft propulsion or production.
power system failure.
However, if the relevant technologies can be
Summary and Conclusions matured, and the systems developed
successfully, the prospective space mission
Without substantial systems-level applications are remarkably broad. In the
innovation and the development of tractable, nearer term, they include Ealth system
but as yet un-demonstrated new transportation, as illustrated in Figure 7;
technologies, a broad range of ambitious while in the longer-term applications include
space operations beyond low Earth orbit cargo (and even perhaps crew)
cannot become either affordable or transportation across the illustration, as
sustainable. illustrated in Figure 8.
Modular, reconfigurable spacecraft, All in all, the MRHE spacecraft concept
assembled in orbit from identical building addresses many highly attractive features in
block components, is a critical capability a robust, scalable package. It promises at
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least one potential solution to an otherwise
insurmountable barrier: how to achieve
affordable and sustainable ambitious future
space operations beyond low Earth orbit.
Glossary of Acronyms
CER Cost Estimating Relationship
EELV Evolved Expendable Launch
Vehicle
ETO Earth-to-Orbit (Transportation)
FTE Full Time Equivalent
GEO Geostationary Earth Orbit
HLLV Heavy-Lift Launch Vehicle
ISS International Space Station
kW kilowatt
kWh kilowatt-hours
LEO Low Earth Orbit
MEO Middle Earth Orbit
MRHE Modular Reconfigurable High
Energy (System)
NASA National Aeronautics and Space
Administration
PDS Power Delivery System
PMAD Power Management and
Distribution
R&D Research and Development
SEP Solar Electric Propulsion
SPG Solar Power Generation
TPI Terrestrial Planet Imager
TRL Technology Readiness Level(s)
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Sun toward viewer.
Figure 1. A Linear / Gravity Gradient Configuration Modular Reconfigurable High
Energy (MRHE) Concept
Robotic Arm
Paytoad
Module Bus (woi:leciric Thruster)
Figure 2. A Tetrahedral / 3-Axis Stabilized Configuration Modular Reconfigurable
High-Energy (MRHE) Concept
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-------------~ ~---- -
Non-power-conducting
roll rings on docking module
(top) and lower side of "pop
out"
Six 2.Sm x Sm SLASR
panels In 2 arrays are
mounted on a "pop-out"
unit offset by a boom from
the SIC. "pop-our unit
rotates at top of boom on a
non-power-conducting
-8m-long booms will
rollring. Two arrays provide avoid plume- LIll-ll.:..-l
25kWe at EOL, can drive 5
Flexible power impingement of
4.SkW thrusters on each
thrusters on arrays:
conducting cable
SIC. deployed with boom, booms do not rotate rrrr-...-....,
with "pop-out" unit
fixed at top and
bottom, winds 360'
around boom during
one orbit, unwinds at
eclipse
Figure 3. A Linear / Normal-to-the-Plane-of-the-Orbit Modular Reconfigurable
High-Energy (MRHE) Concept
Booms do not need to be fully deployed
during build stages; long spacing
between SIC and solar arrays only
needed during thruster operation
Figure 4. Notional Self-Assembly Sequence for a Modular Reconfigurable High
Energy (MRHE) System
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Figure 5 Alternative Large Solar Electric Propulsiojn (SEP) Vehicle Concepts
(Monolithic Square Array (Left) and Thin Film "Bat-Wing" Array (Right)
Preliminary Trajectory TOF Analysis
300 km Initial LEO
-- ISP. 1478 sec: 6.34 N
dfVI COlt I ~~ 9 -- ISP. 1645 sec: 5.70 N
200 -- ISP. Z400 sec: 3.90 N
400 20'2
600 2.c)! -- GEO • 42.100 km
1000 J'B5 a
Results for Continuous Sprial-out
~ 7
Initial Altitude = 300 km
:5
w~ 6 Initial Inclination = 28.45 deg
oE Initial Mass = 13.215 kg
~ 5
'x
«
.Q. 4
'"
!i=
.~ 3 Disregard this
en curve
2
TOF - 112 days to get to an altitude of
-43,600 km (I.e., 50,000 km distance from
Earth center) with Isp=1645 sec (200 V).
Oo~ __- 2L0 _ ___4~ 0 __- 6L0 _ ___e~ o __- 1L00_ ___l~ Z0 __~ 140 _ ___1~ 60 __~ lao _ _~ zo o
Flight Time (days)
Figure 6 Preliminary Flight Times Parametric Analysis for Large SEP Systems (with
varying output voltage)
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