NASA Technical Reports Server (NTRS) 20070002633: In-Space Cryogenic Propellant Depot (ISCPD) Architecture Definitions and Systems Studies

4.7 MB·English

by  NASA Technical Reports Server (NTRS)

Mr. John Fikes
National Aeronautics and Space Administration (NASA)/Marshall Space Flight Center,
Huntsville, Alabama , United States, [email protected]
Mr. Joe T. Howell
National Aeronautics and Space Administration (NASA)/Marshall Space Flight Center,
Huntsville, Alabama, United States, ioe.howellOnasa.gov
Mr. Mark Henley
The Boeing Company, Canoga Park, California, United States,
mark.w. [email protected]
ABSTRACT
The objectives of the ISCPD Architecture Definitions and Systems Studies were to determine
high leverage propellant depot architecture concepts, system configuration trades, and related
technologies to enable more ambitious and affordable human and robotic exploration of the
Earth Neighborhood and beyond. This activity identified ‘architectures and concepts that
preposition and store propellants in space for exploration and commercial space activities,
consistent with Exploration Systems Research and Technology (ESR&T) objectives.
Commonalities across mission scenarios for these architecture definitions, depot concepts,
technologies, and operations were identified that also best satisfy the Vision of Space
Exploration. Trade studies were conducted, technology development needs identified and
assessments performed to drive out the roadmap for obtaining an in-space cryogenic propellant
depot capability.
The Boeing Company supported the NASA Marshall Space Flight Center (MSFC) by conducting
this Depot System Architecture Development Study. The primary objectives of this depot
architecture study were: (I)de termine high leverage propellant depot concepts and related
technologies; (2) identify commonalities across mission scenarios of depot concepts,
technologies, and operations; (3) determine the best depot concepts and key technology
requirements and (4) identify technology development needs including definition of ground and
space test article requirements.
1
INTRODUCTION AND BACKGROUND recommended use of cryogenic hydrogen
and oxygen on two upper stages for Shuttle-
An in-space cryogenic propellant depot
derived launch vehicles as well as a lunar
capability represents a key element of Lander. While the final architecture may
NASA's visions. The servicing of propellants
vary, these are good candidate vehicles for
and consumables in space enables a
ISCPD refueling, and Table 1 summarizes
multitude of mission scenarios, otherwise their expected quantities of cryogenic
unavailable due to costs or operational
propellant. In addition, ESAS recommended
constraints and/or inefficiencies. Cryogenic use of cryogenic methane and oxygen on
Fluid Management (CFM) technology
an Ascent Stage, including storage of these
applications are particularly suited in
cryogens for at least two weeks in space.
evolving capabilities for commercialization ESAS did not specifically require use of an
and solar system science missions. These
ISCPD, but the NASA Administrator
applications cut across all human recommended commercial ISCPD
exploration missions, including depots,
development, with a market value of 2.5
orbital transfer vehicles such as the Crew
billion dollars per year to refuel the ESAS
Exploration Vehicle (CEV) and other in-
Earth Departure Stage (EDS), thereby
space stages. increasing lunar payload mass, and
providing even greater value for Mars
At the core of the depot capability is the
exploration.
economic management of cryogens without
undue or complicated impositions on
Table I. Cryogenic ESAS vehicles could refuel or
infrastructure, other systems, or mission
offload residuals at an ISCPD.
operations. This technology leads to
autonomous fluid management operations
without the complications of propellant
settling and without extravehicular activity
(EVA) support. The basic goal is to enable
automated zero-g storage and transfer of
cryogenic fluids from supply tanks (Figure 1) metric tonnes Stage Stage Lander
to user tanks: safely, reliably, and with
Dry Mass 19 17 6
minimum loss of propellant.
Such vehicles may be refueled at a depot
or deliver residual propellants (to re-fill
depot tanks) and might even store
cryogenic propellant in space. Figure 2
illustrates these vehicles in the
recommended ESAS scenario and how they
may fit lSCPD applications.
Artist Concept.
ARCHITECTURE DEFINITIONS
Recommendations for ISCPD Architectures
were developed based on prior studies and
NASA's concurrent Exploration Systems
Architecture Study (ESAS). ESAS
2
power generation from Earth launch until
lunar ascent, with oxygen reactant stored in
the oxygen propellant tanks, while hydrogen
reactant is stored in the hydrogen propellant
tanks. In this way, the Lander serves as an
ISCPD and it could continue to store
cryogens after the humans return to Earth,
supplying reactants to refuel rovers using
fuel cell power, and receiving and storing
oxygen and hydrogen produced from
resources found on the lunar surface (in
lunar regolith andlor polar ice).
offers opporhrnities for ISCPD uses.
Oxygen and hydrogen production from
In the ESAS-recommended lunar water has the potential to provide great
architecture, the EDS would use roughly benefits for Space Exploration. Water may
123,000 kg of cryogenic propellants to be launched from Earth at low cost or found
reach Low Earth Orbit (LEO) where it would on Earth’s moon, and the moons and
mate with the CEV and burn its remaining surface of Mars. Water electrolysis systems
propellant (-85,000 kg) for Trans-Lunar have been developed for the International
Injection (TLI). If the EDS were refueled in Space Station (ISS) and one was launched
LEO as Dr. Griffin suggested, it could more on the last Space Shuttle mission, to
than double the payload to TLI. A refueled produce gaseous oxygen (for breathing)
EDS could also potentially carry payload and hydrogen (vented overboard). With
into lunar orbit (or initiate Lander descent). refrigeration and heat exchange, we can
Such scenarios could double or triple the convert such gases into cryogenic liquids,
payload mass to the lunar surface. and store them at In-Space Cryogenic
Propellant Depots.
The Cryogenic Upper Stage for CEV
launches is planned for suborbital flight, with The exploration architecture is expected to
re-entry over the Pacific Ocean, but it may grow in capabilities over time. Initially,
also be able to reach orbit carrying the CEV. ISCPD cryogenic propellant may be
In orbit, the remaining tons of cryogenic delivered from Earth, using chemical
propellants could provide low thrust propulsion to reach a depot in LEO. Launch
propulsion, fuel cell power, and systems may deliver cryogenic propellants
Environmental Control and Life Support in dedicated launches as well as by
System (ECLSS) consumables to support scavenging of residual and reserve
the CEV. If this upper stage is carried to propellants from cryogenic upper stages.
orbit, it could off-load residual propellants at Initial ISCPD capabilities may be limited to
an ISCPD, or be refueled and re-used. passive storage (with no refrigeration, but
using boil-off gas to provide propulsion,
The lunar Lander will be designed for power and water), however the architecture
longer-term storage and management of will evolve, with increasing power
cryogenic fluids: roughly one week for requirements to allow for zero boil-off
propulsive maneuvers and two weeks for (refrigeration), and eventually allow for
cryogenic propellant production from water
fuel cell reactants. The Lander’s cryogenic
propellant storage system must withstand launched to the depot from Earth or carried
daytime lunar surface heating, and the from extraterrestrial sources using
Lander design includes radiators for heat advanced propulsion and/or aero-braking.
rejection. Fuel cells on the Lander provide
3
The amount of energy needed to launch selection of oxygen and methane
mass from the moon to LEO or L-1 is much propellants for lunar ascent, as a precursor
less than that to launch from Earth, so it to use for Mars ascent.
may eventually be economical to use lunar
NFIGU
resources to make cryogenic propellants for
use in cis-lunar space and Trans-Mars
ISCPD capabilities are expected to evolve
Injection (TMI). One potential approach
over time, starting with relatively simple
would carry cryogenic propellant from
initial systems, and improving upon these as
production facilities on the moon to an
technologies mature and confidence grows.
ISCPD at L-I, Low Lunar Orbit (LLO) or
For example, an initial ISCPD configuration
LEO. In this scenario, cryogenic oxygen
may use passive storage of modest
and hydrogen are made from ice found in
quantities of propellant in LEO, to serve
cold, permanently shadowed areas near the
human lunar exploration systems, with
moon’s poles. A lunar Lander is refueled
growth to use active refrigeration and store
and launched from the moon’s surface to
very large quantities of propellant for human
low lunar orbit, with further propulsion to
missions to Mars. To the extent practical,
reach L-1 or a trans-earth injection
ISCPD systems should be designed for pre-
trajectory (TEI). From TEI, multiple pass
planned product improvement, with
aero-braking could gradually lower the
configurations allowing a wide range of
perigee to reach LEO without an aerobrake
applications. Initial ISCPD facilities may
shield (a technique previously used at
operate in a micro-gravity environment in
Venus and Mars). In the distant future, a
LEO, with additional facilities emplaced later
more advanced approach might produce
on the moon, at Earth-Moon or Earth-Sun
small “vehicles” filled with water from lunar
libration points, and in lunar orbit. As the
resources; launch them via propellantless
architecture evolves to include In-Situ
rail-gun to reach L-1 or TEI, and then
Propellant Production, depots may also
convert the water into cryogenic propellants
operate in a high gravity environment on the
to be stored in an ISCPD.
Moon, Mars, and the moons of Mars. Table
2 summarizes the expected order of priority
Human missions to Mars will use larger
for ISCPD applications and their different
quantities of propellant, requiring significant
environments.
growth in ISCPD propellant capacity. The
payload sent to Mars may also include
significant amounts of liquid hydrogen (e.g.,
18,700 kg), thus the payload itself may
include an ISCPD for cryogen storage
throughout the long journey, with continuing
storage in orbit around Mars and on the
surface. On Mars, hydrogen from Earth
may be combined with carbon dioxide from
the atmosphere to make cryogenic methane
and oxygen propellants for return to Earth
(2H2+C0p>CH4+02 via the Sabatier
process and electrolysis), along with excess
oxygen for breathing, and water and power
from fuel cells (02+2H2=>2H20+Power).
Such an In-Situ Propellant Production
(ISPP) strategy may significantly reduce the
mass launched from Earth and the cost of
the associated Mars Exploration program,
and this scenario influenced the ESAS
4
Table 2. Order of priority for lSCPD applications
and their different environments.
Exposure
24 hr day;
to 60% of Near 100% on -60-loo%, Near 100%
I Near 45% Of dust-storms; Year-
the time polar mountain bi-weekly (occasional 1.5 solar constant at long day night at
solar & 0% in crater variation hr eclipse) Earth Dnln
I
Heat exchange Radiate to deep Radiate to deep Radiate to deep Dust issues;
with lunar ice? space space space Clouds
Micro- flux*g ravity Large increase NO ‘cup” flux, Naturadle ep Increased flux of Protected by
meteroid & from lunar gravity increase space flux
meteoroids
Orbital meteor ejecta Lunar ejecta atmosphere
Orbital debris No orbital debris
Debris (top & sides) impacts front absent
front &
ISCPD tanks are expected to be quite large. It would be possible to place a very large
We considered using extremely-large tanks, ISCPD in orbit with a single launch. As is
launched as a monolithic structure, as well shown in Figure 3, a depot could be created
as moderately-large tank modules, joined to with a capacity for 400 tons of cryogenic
other modules in space to create a depot. propellant by using Delta IV ELV “Common
Small tanks were only considered briefly, as Booster Core” (CBC) tank-sets in both the
their mass is higher for a given quantity of launch vehicle and as the payload
propellant (due to a higher surface area to (replacing the fairing). During launch,
volume ratio), however small tanks may be propellant is transferred from the upper
needed for high pressure (supercritical) CBC task-set to the lower tank-set, and the
fluid. The main issue for larger tanks is engine burns longer with this added
access to orbit, and the favored design propellant, to place the entire monolithic
solutions are to launch ISCPD tanks as structure into orbit. Such large tank-sets
upper stages using only some of their allow simple “gravity gradient” settling of
cryogenic propellants to reach LEO, and their cryogenic propellants, as the related
continuing to store propellants in orbit. The forces are much less than forces of surface
ISCPD tanks were based on Delta tension in large tanks. phe “Bond number”
Expendable Launch Vehicle (ELV) stages, is very large, Bo = Bond Number =
but the same general logic could apply to (2paR)/o, where a = acceleration, R = tank
other launch systems, including variants of radius, p = density, and cs = surface
the ESAS architecture. tension]. Gravity gradient orientation,
however, is undesirable for atmospheric
5
drag and debris impact hazards, and other
techniques tend to settle propellant to the
- -
same end of the depot (vs. opposite ends). Payload Fairing ----T
*my( n shotler
The large monolithic depot would also
require a single dedicated launch, adding a cry0 upper sta
lnntal ISCPD Configuration,
ISCPD Module
risk of losing the entire depot in a single
.Reach Orbit nearly Full
failure, and it would arrive in LEO nearly or Cryogenic hopeiianf
RCS Circularization .
empty. .Main Engine deleted
*Lwlhrusla dequate
Figure 4. Modular depot tanks can launch as
upper stages, reaching LEO nearly full.
A wide range of techniques could settle
cryogenic propellants for acquisition and
transfer in zero gravity. We expect ISCPD
settling techniques to also evolve with time:
initial settling could use boil-off gas from a
receiving vehicle’s hydrogen tank as a
Figure 3. A large monolithic depot could reach
propellant to provide a low thrust.
LEO with a single launch with near-empty tanks.
Techniques without propulsive thrust will be
required when ISCPD capabilities grow to
More modestly sized depot tanks could
include “zero-vent fill” (with more power and
reach LEO nearly full of propellants,
refrigeration). Techniques include tank
requiring only a small propulsive maneuver
exchange, use of gravity gradient forces,
for orbit circularization after release in a
surface tension, and system rotation or fluid
sub-orbital trajectory. Figure 4 illustrates
rotation (in tanks). Of these, surface
such a depot tank-set launched in place of a
tension systems appear most promising as
cryogenic upper stage. In this scenario, a
a baseline. Another advanced technique
main engine, typically required for upper
could use magnetic fields: since liquid
stages, is not needed, as lower thrust H2-
oxygen is paramagnetic (attracted to a
02 thrusters are sufficient to perform a
magnetic field) and liquid hydrogen is
circularization burn over a long time interval
diamagnetic (repelled by a magnetic field)
at apogee. A single launch provides initial
(note that the Earth’s magnetic field may
depot capabilities, including the delivery of
even need to be considered as an influence
propellants. The configuration can grow
on propellant behavior in LEO).
with the modular addition of more tank-sets,
as well as additional power and thermal
ALTERNATIVE CONCEPT DEFINITION
radiation systems for refrigeration and zero
AND ASSESSMENT
boil-off. The modular approach allows
tailoring of depot propellant capacity to meet Notional ISCPD system configurations were
re-supply needs that change depending defined for comparison purposes and
upon time and depot location with improved alternative conceptual designs also
debris protection. While configuration details assessed. A reference depot module was
may vary, such a modular approach is defined, as summarized in Figure 5. The
recommended as the most practical course module uses a thermodynamic vent system
for gradual development of ISCPD for hydrogen boil-off, with H2 gas passing
capabilities. through a vapor cooled shield on the tank
6
wall, then conducting heat away from the orientation exposes different parts of the
oxygen tank before venting. A contingency module to different environments; LEO
vent system is included in the oxygen tank. orbital debris hazards are most severe from
The module shown in Figure 5 has a the sides, meteoroids and sunlight come
deployable multi-layer insulation blanket, from above, and the Earth’s heating
which also provides protection for the (infrared and albedo) comes from below,
hydrogen tank against micrometeoroids and thus modules surface details (insulation,
orbital debris. Rigid insulation alternatives thermal shielding, etc.) may be tailored to
also have merit. Pressurization is best meet these differing conditions.
autogenous, using small tanks of
supercritical H2 and 02 gas, which also
provide fuel cell reactants and RCS
propellants. The module includes
accommodations for autonomous docking
and fluid transfer on both the forward and
aft ends.
Figure 6. Growth ISCPD Facility in LEO: Add
propellant and power for refrigeration.
Autogenous pressurization is important for
the depot (and stages that it refuels) to
avoid requirements for re-suppling high
pressure helium gas, which is difficult to
contain and transfer. Cryogenic liquid is
transferred to a small “boiler tank where it
is warmed using thermal switches and heat
exchangers to reach high pressure,
becoming a supercritical fluid. This warmer
Figure 5. Reference ISCPD module concept.
fluid is then transferred to a Composite
Overwrapped Pressure Vessel, and which
in turn supplies H2/02 gas-gas Reaction
Control System (RCS), fuel cells, and H2
Figure 6 illustrates growth of the ISCPD with low thrust propulsion systems as well as
additional modules and solar power. Solar providing pressure for the cryogenic liquid
power as shown, is based on existing tanks (to force fluid to transfer from the
satellite solar power systems, and is sized depot into lower pressure tanks on the
for a 20 kWe peak power level (roughly 10 receiving vehicle).
kWe average in LEO). In this view, one can
see a preferred orientation with respect to TECHNOLOGY DEVELOPMENT
the Earth. This orbital orientation minimizes PLANS
drag, which tends to settle propellants
forward, in the direction of the orbital Plans for development of critical ISCPD
velocity vector (equivalent to “downward” in technologies include a potential space flight
the launch orientation). The low drag demonstration program and ground
7
demonstration options that prepare for
flight-testing. Cryogenic fluid management
in zero- or micro-gravity has been analyzed
extensively with few opportunities to verify
analytical models in space. The Apollo-
Saturn 203 Flight was dedicated as an
experiment to monitor cryogenic propellant
conditions and dynamics in orbit; however
this approach is fairly costly. Relevant
space flight data can also be gained without
significant cost, however, when flight
experiments are performed as a secondary
Figure 7. Cryogenic upper stages vent hundreds
mission objective on a cryogenic upper of pounds of remaining fluids that could be used
stage, using its remaining cryogenic for secondary flight experiments on virtually every
propellants after the primary payload is mission.
released. The Titan-Cenatur-2 Mission used
this approach to perform two additional After the spacecraft deployment occurs and
firings of the engine after storing cryogenic the spacecraft reaches an acceptable
propellants for I-hour and 3-hour coast distance from the Cryogenic Upper Stage,
intervals, and to demonstrate a “bubbler” the stage reorients to a new position. The
system to reduce helium usage (by CCAM moves the stage away from the
increasing the oxygen partial pressure). spacecraft orbit to prevent collision, and
The Titan-Centaur-5 launch of Helios-2 also expels propellant (in a direction away from
used this approach to demonstrate a total of the spacecraft) to increase the separation
seven burns of an RL-10 engine in a variety distance and relative velocity, and to
of conditions and storing cryogenic prevent subsequent tank rupture. Today’s
propellants for 5.25 hours between burns. cryogenic upper stages typically complete
Many future NASA launches could use a their primary missions with significant
similar strategy to experiment with masses of leftover fluids (hundreds of
cryogenic propellants remaining in kilograms), including cryogenic liquids
expended upper stages after their primary (residuals, reserves, trapped fluids, and a
payloads are released. Typical cryogenic hydrogen bias), cold gas (hydrogen ullage
upper stage mission event sequences gas, oxygen ullage gas, and residual
deploy the payload(s), then perform helium), and even some hydrazine RCS
contamination and collision avoidance propellant.
maneuvers (CCAM), including venting of
remaining cryogenic fluids into space, as is
illustrated in Figure 7.
8
CONCLUSIONS
In-Space Cryogenic Propellant Depot
systems offer significant advantages for
NASA space exploration systems.
Refueling of in-space transfer stages at an
Reorient Secund &age Perform CCAM
for Contamination end to Remove ISCPD can support NASA’s ESAS lunar
Collision Avddallce SemndStqefrom
Maneuver (CCAM) Spacecraft Orbit exploration architecture and may be
enabling for human exploration of Mars.
Upper Stage Reorients and Secondary Plight
Performs Collision Experiment Uslng Plnid ISCPD sizing is expected to be moderate,
Avoidance Maneuver (not Residuals & Reserves
Spacecraft propellant venting) (HZ, 02, He & N2B4) allowing deliver of modules to LEO as upper
Separation
stages without main engines, nearly full of
propellant. ISCPD design recommendations
Figure 8. Secondary flight experiments on include modular construction and features
cryogenic upper stages may use residual
allowing autogenous pressurization (without
propellants to test technology instead of dumping
helium gas). Technology demonstrations
them shortly after payload separation.
may use secondary experiments on
As shown in Figure 8, simple flight
cryogenic upper stages as a means for
experiments may test maneuvers or new
ready access to orbit.
hardware after the primary mission (e.g., to
settle propellant or gauge its mass), and
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