IEEE - Aerospace and Electronic Systems - September 2019 - 16

Water Electrolysis Propulsion as a Case Study in Resource-Based Spacecraft Architecture (February 2020)
If propellant is among the resources being combined by
the RSS method, this approach becomes recursive. It also
becomes more complex if multiple stages are used.

EXAMPLE: THE WORLD IS NOT ENOUGH
This section considers the world is not enough (WINE)
mission concept, a steam-propelled design which collects
water in situ to replenish propellant supplies as part of a
sustained campaign of exploration of asteroids, icy moons,
or other suitable Solar System bodies [28]. We consider
the application of electrolysis propulsion as a higher efficiency addition to or replacement for the steam thruster
used in the WINE design, an idea suggested by the WINE
team themselves.
As mentioned in the previous section, the WINE mission concept is a collaboration between the University of
Central Florida, Embry-Riddle Aeronautical University,
and Honeybee Robotics. One disadvantage of the WINE
steam engine is a low specific impulse. The design used for
WINE has a specific impulse of 160 s [33], and even
nuclear thermal steam propulsion designs have a low specific impulse of less than 200 s [20]. This is significantly
better than cold gas thrusters, but significantly worse than
demonstrated electrolysis propulsion performance of close
to 300 s and the potential for performance as high as 450 s,
close to the ideal for chemical propulsion [5], [17]. The relative impulse density of steam propulsion is therefore
approximately 8-10, compared to the 15 of present electrolysis propulsion performance, and the 23 of ideal electrolysis propulsion performance.
The WINE team recognizes this. They suggest electrolysis propulsion as a possible addition to the mission
concept that would have better performance for the relatively high DV transfers between bodies [28]. The 6U version of WINE proposed for asteroid prospecting has a
mass of 7 kg and baselines a 3U volume for the propellant
tank. A full propellant load of 3 kg would therefore produce 560 m/s DV using steam propulsion. Electrolysis
propulsion almost doubles the available DV per propellant
load, providing 1050 m/s DV at present day electrolysis
propulsion performance of approximately 300 s specific
impulse. At the ideal electrolysis propulsion performance
of 450 s, the available DV almost triples, to 1574 m/s per
propellant load.

EXAMPLE: MARS DESIGN REFERENCE MISSION 5.0
This section considers NASA's Mars Design Reference
Architecture 5.0 (DRA-5) with the key change made that
the propellant is stored as liquid water instead of LH2 or
LH2/LOX, providing commonality between propellant,
oxygen generation stock, and water supplies for the crew
[2]. The DRA-5 study proposes several propulsion
16

options; some use only impulsive maneuvers while others
use low-thrust maneuvers or a combination of both.
Impulsive propulsion concepts considered include nuclear
thermal and cryogenic chemical bipropellant. Low-thrust
concepts considered in the DRA-5 include nuclear electric propulsion and a combination of solar electric and
chemical propulsion. Electrolysis propulsion compares
most favorably to the cryogenic chemical propulsion
option, because it can achieve the same specific impulse
but without the need for long duration cryogenic propellant storage.
The DRA-5 chemical propulsion mission concept
budgets five super-heavy lift launches and 347.6 metric
tons of propellant for the crewed vehicle including margin
[11]. For the purposes of this example, we assume liquid
water is used for propellant storage instead of LH2/LOX
and ignore the possibility for dry mass savings based on the
easier handling of water. For the transit habitat, 2.65 m of
consumables are budgeted for the outgoing trip as well as
for the returning trip, with an additional 7.94 m of contingency in case the crew must remain in orbit and cannot
reach the surface habitat and supplies. In addition to these
consumables, 0.687 m is budgeted for contingency water
and oxygen for crew use. This optimistic design decision is
based on the assumption that "recovery of water from urine
is assumed at 85% whereas all other water sources are
recovered at 100%. . . Water is not required for the transit
or surface habitats because of the amount of water assumed
in the food recovered by the ECLSS" [11]. Based on a
2035 launch window, there is additionally a 31.6 m total
propellant margin across all interplanetary maneuvers:
trans-Mars injection, Mars orbit injection, and trans-Earth
injection. Combined, a total contingency of 32.287 m of
water is carried.
Due to the shared resource of water between propulsion and crew supplies, the total contingency mass
required is reduced using the RSS method described
above. The new result is 31.607 m, saving 0.68 m. This
mass reduction is modest, because the RSS is dominated
by the large propellant term. However, this initial reduction cascades into additional mass savings, because with
the reduced mass carried as contingency, less propellant is
needed to achieve the same DV for each maneuver. We
save an additional 2.68 m of water in this way, for a total
of 3.36 m.
More importantly, technical and logistical gaps in the
DRA-5 may be bridged if water electrolysis propulsion
technology is baselined in place of an extreme duration
propellant-storage solution that currently does not exist.
Both the nuclear thermal and chemical propulsion
options for the DRA-5 require storage of cryogenic fluids
for long periods of time, approximately two years. As
noted in the previous section, the study considers this a
"critical technical area" and the longest endurance of
stored cryogens on-orbit is 9 h [11]. Electrolysis

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