Mars-Base-Camp-Update-and-New-Concepts

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© 2014 Lockheed Martin Corporation

Mars Base Camp
Updates and New Concepts

A Technical Paper Presented by:

Timothy Cichan
Lockheed Martin Space
timothy.cichan@lmco.com

Kerry Timmons
Lockheed Martin Space
scott.d.norris@lmco.com

Kathleen Coderre
Lockheed Martin Space
robert.p.chambers@lmco.com

Willian D. Pratt
Lockheed Martin Space
william.d.pratt@lmco.com


September 2017

Abstract

Orion, the Multi -Purpose Crew Vehicle, is a key piece of the NASA human exploration architecture
for beyond earth orbit. Lockheed Martin was awarded the contract for the design, development,
test, and production for Orion up through Exploration Mission 2 ( EM -2). Additionally, L ockheed
Martin is working on defining the cislunar proving ground mission architecture, in partnership
with NASA , and exploring the definition of Mars missions as the horizon goal to provide input to
the plans for human exploration of the solar system.

In 2016, Lockheed Martin presented a proposal for achieving crewed exploration of Martian space
as early as the 2028 launch opportunity. Known as Mars Base Camp, this proposal involved
establishing a crewed vehicle in Martian orbit fro m which astronauts could perform excursions to
Deimos and Phobos, and could also perform telerobotic exploration of the Martian surface,
including sample return. This concept presented a novel, practical and affordable path to enable
human exploration of the Martian system in the next decade. This paper will detail additional
development for the Mars Base Camp concept, including the production of propellant from water,
additional details for the cislunar proving ground missions, and a Mars lander concept.

The orbiting base camp could generate oxygen and hydrogen from water via solar -powered
electrolysis. Water may be provided directly from the Earth system or via in -situ resource
production in the lunar, Martian, or other systems. The demonstration of Mar s Base Camp
capabilities at the Deep Space Gateway will be discussed, including systems, technology and
scientific mission possibilities. The lander is envisioned as a fully reusable, lifting body that uses
supersonic retro -propulsion to descend and land o n the surface. Initial crewed missions using the
lander, which would follow on later missions than the initial mission, are outlined as relatively
short -duration, science -focused exploration missions.

Multiple areas of the Martian surface would be explored with the objective to gather scientific
data from a wide variety of sites of interest, and more fully characterize possible sites for future
permanent settlements. Once a surface mission is completed, the lander returns to Mars Base
Camp as a single stage to orbit launch vehicle to be refueled. With these additional developments,
the Mars Base Camp concept can be seen as a core system that pivots humanity into a viable,
sustainable long -term Mars exploration program.

Acron yms/Abbreviations


∆V = Velocity Change
CM = Crew Module
CPS = Cryogenic Propulsion Stage
DSG = Deep Space Gateway
DSN = Deep Space Network
ECLS = Environmental Control and Life
Support
EDL = Entry, Descent, and Landing
EVA = Extra Vehicular Activity
Isp = Specific Impulse
ISRU = In-situ resource utilization
IVF = Integrated Vehicle Fluids
L/D = Lift to Drag Ratio
LOX = Liquid Oxygen
LH2 = Liquid Hydrogen
MBC = Mars Base Camp system
MBC -N = The Nth Mars Base Camp Mission
MBC -S = MBC Mission with Surface Sorties

MADV = Mars Ascent/Descent Vehicle
MT = Metric Ton (1000 kg)
OML = Outer Mold Line
RCS = Reaction Control System
SEP = Solar Electric Propulsion
SLS = Spac e Launch System
Sol = One Martian Day (24h, 37m, 23 s)
SRP = Supersonic Retro Propulsion
TPS = Thermal Protection System
TRL = Technology Readiness Level
WDV = Water Delivery Vehicle
1. Introduction

In 2016, Lockheed Martin introduced Mars
Base Camp: a Martian Moon Human
Exploration Architecture 1. This architecture
is designed to enable human exploration of
cis -Martian space on an accelerated
timetable. The first mission utilizing this
architecture, deemed Mars Base Camp 1
(MBC -1) e nables a crew of 6 to spend a year
in Martian orbit performing real -time
telerobotic operation of rovers and
unmanned aerial vehicles on the Martian
surface, and also performing crewed sorties
to the surface of Deimos and Phobos. This
architecture also enables humans to become
an interplanetary species, and would return
a sizeable amount of science value due to the
ability to do simultaneous, real -time crewed
exploration of multiple areas in the Martian
system. However, perhaps the greatest
return on investment f rom the MBC -1
mission is the reusable infrastructure it puts
in place to enable affordable, ongoing
crewed missions to Mars of increasing scope
and complexity.

This paper builds on the Mars Base Camp
architecture proposed and also discusses the
scientific operations and opportunities at the
Deep Space Gateway and how these
technologies and processes enable and
directly translate to operations of Mars Base
Camp. These mission capabilities include
telerobotics, sample return, 3D printing for
spare parts and tools, and solar electric
propulsion.

This paper also establishes the concept of
operations whereby the elements placed in
orbit during MBC -1 can be used to enable
crewed exploration of the Martian surface in
follow -on MBC sortie missions (MBC -S). Core
elements of the MBC transfer vehicle that
remain in high Earth orbit after MBC -1 are
refueled, and used as the transfer vehicle
onto which two Orions are docked for
transfer to Mars for the MBC -S mission. This
time, however, the pre -placed Martian
elements i nclude at least one Mars
Ascent/Descent Vehicle (MADV) and,
optionally, one or more separate, Mars -
orbiting, cryogenic fuel depots. These
additional elements now enable MBC to be
used as the basis of operations from which
crewed sorties to the surface of M ars can be
executed.

Multiple, short -duration sorties to the
Martian surface can be performed during the
ensuing 11 -month stay at Mars (if the
separate orbiting fuel depots are assumed).
This paradigm enables maximum flexibility in
crewed exploration of t he surface; landing
sites can be adjusted in near real -time during
the mission as targets of opportunity are
remotely identified. In addition, the safety of
these early surface missions is greatly
enhanced as the crew can return to MBC at
almost any time i f equipment issues, medical
concerns, etc. dictate. This includes the
ability to abort to orbit at any point during
Mars Entry, Descent and Landing.

Mars Base Camp is not a “flags and
footprints” system. Rather, MBC -based
missions not only enable human ex ploration
of Mars on a near -term timescale, but also
establish an infrastructure of reusable
elements that enable sustainable, long -term
crewed operations at Mars. T his system is
intended to embrace and leverage
internatio nal and commercial partnerships ,
and several opportunities for purely
commercial service as part of the overall
concept of operations are identified.





2. Key Tenets for Long -Term Martian
Exploration Architecture

The purpose of Mars Base Camp is to
establish a vision for a realistic, achievable
architecture whereby crewed exploration of
the surface of Mars may be achieved within
the timeframe (2030s) and budget currently
outlined by US policymakers. Furthermore,
the vision for Martian explorat ion presented
here is driven by several main tenets:
 Each mission should lay the groundwork
for the next: We here embrace the
overall vision that the foremost goal of
Martian human exploration should be to
establish a sustainable long -term human
presence a t Mars. “Flags and footprints”
missions – where a crewed landing has
no viable follow -on without the
development of an entirely new
architecture – lead to large post -mission
gaps and are to be avoided. Each mission
should instead be designed
fundamentally as a stepping stone on a
path to longer surface missions rather
than an end unto itself.
 Reusability = Affordability: The viability
of a long -term Martian exploration policy
depends upon the ability to get as much
use out of the most expensive system
elements as possible. Sacrificial,
expendable modules are to be avoided in
favor of reusable systems where each
element deployed in a given mission is a
key piece of a long term exploration
infrastructure for future missions.
 Leverage existing systems and
technologies to the greatest extent
possible: This was one of the
fundamental tenets of Mars Base Camp
when fi rst established . When such trade
opportunities arise, the architecture
should favor existing and/or higher TRL
technologies that can be deployed o n a
shorter timeframe with lower
development costs over potentially
better performing, but more expensive
Figure 1: Mars Ascent/Descent Vehicle

and longer -lead technologies. The
architecture should be evolvable such
that newer technologies can be
incorporated on future missions as they
become practical. MBC was based
around Orion and SLS, current NASA
habitat development efforts, and current
NASA solar electric propulsion
development efforts for this reason.
 Crew safety and reliability is paramount:
By its nature, interplanetary travel
presents very limited opportunities for
abort and early return. Thus, the
architecture must seek to eliminate
elements or operations where a single
failure could lead to loss of crew, and
should enable opportunities for self -
rescue to the greatest extent possible.
For this reason, most of the elements of
Mars Base Camp are redundant – 2
Orions, 2 crew quarters, potentially 2
MADVs, etc.
 Embrace and foster partnerships
between NASA, international, and
commercial contributors: As with the first
Mars Base Camp Study, this architecture
is envisioned as NASA -led, but with
extensive collaboration / partnering with
international and commercial
contributors. Where possible,
opportunities for certain portions of the
system operations to be realized via a
purely commercially procured service are
to be encouraged and highlighted.

3. Demonstrating Mars Base Camp
Mission Capabilities at the Deep Space
Gateway

At the NASA Advisory Council in May of 2017,
NASA outlined the human space exploration
phases starting from the ISS to t he surface of
Mars. This plan outlines the Deep Space
Gateway buildup with the Habitation module
being delivered on EM -3. The habitation
module enables a crew of 4 to live and work
in cislunar space for at least 30 days. Follow -
on missions, EM -4-6, bring a dditional
Gateway elements and are thought to be
ideal candidates for science missions. The
Deep Space Gateway is an ideal location to
demonstrate many of the mission
capabilities that will be needed for Mars Base
Camp, including telerobotics, sample retur n,
and solar electric propulsion. Just like at
Mars, these capabilities enable significant
science return.

Lockheed Martin is developing an evolvin g
architecture for human space exploration
based on the capabilities of Orion and the
Deep Space Gateway th at will incrementally
push human exploration towards Mars. Mars
Base Camp extends and leverages the
mission operations and technologies
demonstrated at the Deep Space Gateway
for telerobotic operations, sample return,
and system maintenance.

A key piece of the Deep Space Gateway
architecture is to perform robotic operations
and use telerobotic systems to perform
surface science from Orion and the Gateway.
Design concepts are being developed that
include an integrated science and robotic
subsystem which pr ovides flexible interfaces
for control of surface robotics, display of any
data and telemetry , and operational
planning tools. This work includes
developing an understanding of the
limitations and constraints on robotic
capabilities (power, communications,
mo bility, etc.). T elerobotics is extremely
powerful in allowing humans to remotely
operate rovers or UAVs on a planetary
surface, particularly with tasks that are
complex and require quick decisions such as
drilling, driving, and sample collection 2. As an

example, the Curiosity and Opportunity
rovers were operated by ground teams on
Earth, but the tasks were limited based on
power and the round trip communication
time, which range between 380 and 2670
seconds, depending on the Earth -Mars
orientation . With scientist -astronauts
telerobotically operating rovers and UAVs
from orbit, an increased volume of science
can be collected due to the real -time control
of the astronauts as well as the significantly
smaller round -trip light time between the
operator s in orbit and the Martian surface 3.

To continuously operate rovers and UAVs
from orbit over extended periods of time,
the robotic systems need to be designed with
better power storage, a communication
system that supports teleoperation, and the
mobility and manipulation to support
scenarios with astronaut teleoper ation.
Additionally, the robots will need to have
multiple control methods including
telerobotic, automated, human supervised
autonomy , and task level autonomy. Of
course the habitat and laboratory modules
need to support the astronauts performing
telerobo tics by scheduling in operations
within their daily activities to align with
opportunities for robotic surface operations,
provide the astronaut situational awareness
of the robot surroundings through a digital
twin technology fusion, and provide the
astro naut the ability to support any robotic
system from a single robotic workstation.
The process, planning, and methodologies of
scientist -astronaut operating rovers from
the Gateway is directly applicable to the
same scenarios on the Martian surface.
Error! Reference source not found. shows a
scientist astronaut simulating lunar rover
operations from a ground mockup .

In both the NRC -2007 and the most recent
decadal survey report ranked sample returns
from the lunar far side and the Martian
surface as the highest priorities 4. Burns, et al.
have proposed sample returns from the
South Pole -Aitken (SPA) basin and the
Schrö dinger basin 5. The Deep Space Gateway
can support both teleoperated rovers as well
as receiving and partially analyzing sample

Figure 1: Scientist Astronaut Operating Rover on the Moon

returns. This process is applicable to MBC, as
astronauts can perform the same planning
and sample return collection and analysis . In
a Martian orbit, the initial analysis by a
scientist -astronaut is invaluable in assessing
the sample and planning the next one as
opposed to waiting for the sample to return
to Earth for analysis and further planning.
Sample collection capabilities ar e already
being planned into the Mars2020 rover. The
addition of a sample launcher element would
allow collection of the sample in a Mars
orbit, see Figure 2 for a visual representation.

The Deep Space Gateway allows for
astronauts to live and work in space for 30 -
60 days with increasing durations. These long
durations and the remote location of
Gateway provides the astronauts the
opportunity to maintain and operate the
vehicle without the fallback of a quick Earth -
return as available at ISS should an issue
arise. The Deep Space Gateway platform
allows for the development of 3D printers
and the processes for printing and installing
spare parts as opposed to bringing spares.
On 1,000 day mis sions to Mars it is unrealistic
to bring tools and spare parts for every
contingency situation.


Figure 2: MBC Capture of Mars Sample Return
One of the key developments in deep space
exploration, both in going to the Moon and
Mars, is a more efficient means of
propulsion. Solar Electric Propulsion (SEP)
uses electricity generated from solar arrays
to ionize atoms of the inert gas xenon. This
sy stem is low in thrust, but is extremely
efficient with a high specific impulse, I sp. Until
recently, solar electric propulsion has been
used primarily for low power applications
such as station keeping. The Power
Propulsion Element (PPE) of the Deep Space
Gateway, will use SEP to reach its initial orbit
in cislu nar space. After mating with the
habitat element, the PPE will be used to
move the uncrewed habitat and additional
Gateway elements to various orbits around
the moon.

To achieve these goals , the req uired power is
at least 40kW, which is many times larger
than thrusters currently flying, but is scalable
to higher power systems with advancements
in solar array technology and lab
developments of thrusters and power
processing units. Development and test ing of
higher class power systems at the Gateway
enables the exploration of Mars. Power
propulsion systems using SEP, similar to the
PPE, would be used to preposition Mars
exploration elements, like supplies, landers,
and EVA modules, as seen in Figure 3.


Figure 3: Deep Space Gateway Elements Enabled
by Solar Electric Propulsion

4. MBC -S: Con cept of O peration s for
Humanity’s First Steps on Mars

Preparation for the MBC -S mission begins
immediately upon the conclusion of the
previous MBC mission. All system elements
that were fielded in the MBC -1 mission
(habitat and laboratory modules, LOX/LH2
tank farms, crew quarters modules,
cryogenic propulsion stages, solar power
generation stages) remain in space and are
used again to provide Earth/Mars transit and
the living environment in Martian orbit
during MBC -S. This time, however, a Mars
Ascent/Descent Vehicle (MADV) is also pre -
deployed to Martian orbit and docked t o the
MBC station.

This MADV is then used to enact relatively
short, science -focused sorties from MBC to
sites of interest on the Martian surface.
Propellant consumed by the MADV for
surface sorties is generated in Mars orbit via
electrolysis performed on water that is
supplied, ideally, by commercial providers.
Initial missions assume water would likely be
provided from Earth. However, commercial
development of in -situ water sources
(asteroids, Mars surface, and/or ideally on
the moons of Mars) are encour aged as this
represents an opportunity for significant
savings in the recurring cost of a sustained
exploration and development effort.

The MBC -S mission is divided into five
primary phases:
A. Preparation of existing MBC elements for
Earth/Mars Transit duri ng MBC -S. This
includes:
a. Capture of the MBC Earth/Mars
transit elements into orbit with the
Deep Space Gateway.
b. Refueling of the two Cr yogenic
Propulsion Stages (CPS) and tank
farms.
B. Preparation of pre -staged elements at
Mars. This includes:
a. As -needed adju stment of the orbit on
the MBC elements that remain at
Mars (Laboratory module, center
node, Deimos/Phobos excursion
vehicle) into the operational orbit for
MBC -S.
b. Launch of MADVs and rendezvous
with existing MBC elements at Mars.
c. Launch of fuel depot and rendezvous
with existing MBC elements.
C. Launch of crew and transfer to Mars on
the MBC transfer vehicle
a. Same general concept as MBC -1.
b. Final assembly of the full MBC at
Mars.
D. Mars Surface Sorties
a. Fueling of MADV.
b. Descent and Landing.
c. Surface operations.
d. Launch and Return to MBC.
E. Return to Earth
a. Same general concept as MBC -1.
This overall mis sion concept is illustrated
in Error! Reference source not found. .

At the conclusion of the previous MBC
mission, the crew returns to Earth in a subset
of the MBC system called the Transfer
Vehicle. The Transfer V ehicle approaches
Earth on a hyperbolic trajectory and
performs a braking burn to return it back to
the cis -Lunar orbit with the Deep Space
Gateway (DSG) where it originated its
journey. It should be noted that, while return
from the DSG is baselined, Orio n has been
sized for entry velocities up to 11.5 km/s and
thus direct entry with Orion from a
Mars/Earth transit orbit is possible in a
contingency. Once the Transfer Vehicle has
been returned to the Deep Space Gateway, it
is ready to be refueled. Fueling the Tank
Farms in preparation for the mission is a
potential commercially procured service.
Standard propellant transfer interfaces will
be defined and implemented on the CPS and
Tank Farms, providing an opportunity for
delivery of these propellants to the Transfer
Vehicle by multiple commercial and/or
international partners. The Mars Base Camp
Tank Farms are zero boil off designs that
include active cryo genic cooling. The Cryo
Stages are not. Therefore, transfer of
propellants to the Cryo Stages happens ju st
prior to the launch of the crew.

A portion of the MBC system remains in Mars
orbit following MBC -1, as seen in Figure 4:
 Lab Module
 Center Node
 Deimos/Phobos Excursion Module
 2 Solar Electric Propulsion Stages and
related solar panels

Beginning on MBC -S, a new element is added
to the system to enable sorties to the
Martian Surface: The Mars Ascent/Descent
Vehicle (MADV). The detailed concept of
operations for the use of the lander is
described in phase D below. In phase B, the
lander is pre -staged at Mars in preparation
for later a rrival of the crew (see Error!
Reference source not found. for an
illustration of all mission phases).
Figure 5. MBC -S Overall Mission Concept

One or two MADVs are launched – each on
an SLS Block IB.

While the entire MBC -S mission could be
performed with only one lander, use of two
MADVs is highlighted as an opportunity in
this study to enable self -rescue in the event
that the crew becomes stranded w hile on the
surface. A MADV is launched empty, with the
attached dual mode SEP Stage providing
services including rendezvous and docking to
MBC. (Dual -mode SEP stage provides both
solar electric propulsion for long -period
velocity change maneuvers and chem ical
thrust for attitude control and small
translational maneuvers during rendezvous
and docking). The SLS provides injection to
Mars, and the SEP stage then controls
attitude and trajectory as it captures at Mars
and spirals into the chosen MBC -S orbit. W ith
the MADV(s) pre -positioned at Mars, the
system is ready for the crew to arrive.


Figure 5 – MADV Launch Configuration atop
SLS IB
4.1 Mars Surface Sorties
The Mars Base Cam p system presents a new
and flexible approach to exploration of the
Red Planet. The crew will have to be at Mars
for around 11 months in any mission where
Hoh nman n transfers are used for the
Earth/Mars transits (driven by the synodic
period of the Earth/Mars system). In an MBC -
based mission, however, the crew need not
be on the surface, or even at one location on
the surface, during this entire timeframe. In
the MBC paradigm, initial landings on Mars
are envisioned as relatively short -duration
science expeditions. A crew of 4 descend s to
the surface to spend up to two weeks in a
location collecting data and samples related
to key science objectives. They then return to
MBC to perform more thorough analysis on
the collected samples in the MBC Lab
Module and, ideally, refuel the lander for
another sortie.

Figure 4 - Elements Remaining in Mars Orbit
following MBC -1

This approach offers several key advantages
over a more traditional Mars exploration
architecture where the crew spends the full
duration of the mission in a base on the
surface:
 Cost: A year -long stay on the surface of
Mars requires additional system
elements (habitat modules, power
generation systems, rovers, provisions,
tools, etc.) to be pre -positioned on the
surface ahead of the crew. T hese systems
(and the related cargo delivery systems
to get them to the surface) represent
development, procurement, and launch
costs that are not needed for the first
mission if the surface stay is of a duration
where all provisions needed are
contained w ithin the lander itself.
 Safety: Short surface sorties from an
orbiting base of operations
fundamentally reduce or eliminate
several failure modes that would result in
loss -of-crew in a traditional architecture:
1) Reduced dependence on landing
accuracy: A year -long surface stay
requires the crew to land in close
proximity to the pre -positioned
surface elements. In the MBC
paradigm, this mode of failure is
eliminated entirely.

Figure 7 - Mars surface sortie sequence
2) Any time a bort capability: While MBC
sorties are designed to last around a
week, the lander can in fact return to
the orbiting base camp early if
warranted by equipment failure,
medical emergencies, etc. Crew can
also abort to orbit during a failed
landing attempt.
3) Self -rescue capability: If two landers
are assumed, the crew that remains
at MBC during the sortie has the
ability to rescue the crew on the
surface in the event that they
become stranded.
 Flexibility: The landing site can be chosen
and adjusted in near real -time. A sortie
can target any area on the surface that is
accessible from the operational orbit of
MBC (which is the entire surface of Mars
if a polar orbit is assumed). If capability to
refuel the lander in -orbit between sorties
is assumed, then the crew can actually
visit multiple sites, separated by large
Figure 6. Mars Base Camp – Complete MBC -S
Configuration with MADVs

distances, in a single mission. This not
only has the potential to greatly increase
the net science return for a single MBC
mission, but also allows crews to explore
a wide variety of candidate sites for a
future surface outpost before
committing to a single location.

All propulsion, attitude control and power
generation systems on the MADV are
designed to utilize LOX/LH2. Consumable
water for the sortie is designed to be
recovered from power generation fed by
cryogenic boil -off. Thus, LOX and LH2 are the
only resupply commodities that are required
in order for the lander to execute additional
sorties. The number of sorties that may be
executed in a single MBC mission is therefore
limited only by the quantity of LOX/LH2 that
is available in Mars orbit.

The cryogenic tank farms on MBC were sized
to provide propellant for the round trip to
Mars as well as two sortie missions, first to
Phobos and later to Deimos using the Orions,
the Excursion Systems and the Cryo genic
Propulsion Stages. If no Phobos/Deimos
excursions are assumed for MBC -S, then the
MBC system actually holds enough
propellant at the time of Earth departure to
fuel one MADV sortie without any resupply.

Performing more than one MADV sortie
during a MBC mission requires additional
LOX/LH2 beyond what can be carried
onboard MBC. In the MBC -based exploration
paradigm, resupply cryogenics are
envisioned to be generated onsite by
splitting water via electrolysis. The delivery
of large quantities of water to M BC in orbit
therefore presents a major enabler for long -
term sustainment of an MBC -based
exploration architecture, and a significant
opportunity for development of a
commercial industry to provide the water.

In all refueling scenarios, water is delivered
to MBC via an autonomous Water Delivery
Vehicle (WDV). WDVs could theoretically
come in any size, but for the purpose of this
study we assume a unit with 52 MT water
capacity (2 WDVs required to refuel 1 MADV)
and a 375kW -class solar powered electrolysis
plant that doubles as a SEP stage. The WDV
is a self -contained fuel station and consists
of:
 Tankage for 52MT of water (full when
launched)
 Tankage for 40MT of LOX/LH2 (6:1 ratio.
Empty when launched)
 A 375kW -class SEP stage (375kW at Earth
= 160kW at Mars )
 Electrolysis system (shares common
power generation system with the SEP
stage).
 Navigation and Communications systems

A WDV with this configuration will hereafter
be referred to as a 50 MT Class WDV. Initially,
WDVs would be launched from Earth. The 50
MT class WDV pictured above would require
an SLS, but delivery of smaller variants is
possible and could open the possibility for
delivery of resupply consumables via a
competitive commercial market. Over the
long -term, water provider companies could
ultim ately develop in -space water
production facilities to refill WDVs anywhere
water ice and/or hydrated minerals are
present (polar deposits on Earth’s moon,
asteroids, Martian surface, etc.).

Propellant remains in water form during the
transit to eliminate boil -off loss (while MBC
is designed to be zero boil -off, the MADV and
WDVs are not) and to ensure that the full
power output of the solar arrays is utilized to
power SEP (full power output o f the arrays is
needed in both SEP and electrolysis modes).
Once in Mars orbit, each WDV performs a
rendezvous with MBC, is captured by one of
MBC’s robotic arms, and begins converting
its water load into LOX/LH2. Two 50 MT class
WDVs working in parallel w ould create the
fuel for one MADV sortie in about 2.5
months.

The number of sorties that may be
performed during a single MBC mission is
primarily driven by the number of WDVs that
are present. If more than 2 WDVs are
planned, space limitations at MBC wi ll likely
mean that it is more practical to have the
WDVs orbiting separately as a sort of fuel
“depot.” The separately -orbiting WDVs
would be located in the same orbit as MBC,
but phased slightly ahead or behind. Rather
than filling up with cryogenics at MBC, a
sortie would begin with the MADV
performing an autonomous phasing
maneuver using a small amount of
propellant drawn from MBC to rendezvous
with a WDV.

In any of the above options, once fueling is
complete, the crew loads supplies for the
sortie, t hen undocks and performs the
descent/landing sequence. 4 crew go to the
surface while 2 remain with MBC.

4.1.1 Descent and Landing
Once undocked from MBC, the MADV
performs a deorbit burn and proceeds to
perform direct Mars atmospheric entry. The
analysi s presented in this paper has sized the
MADV for entry velocities up to 4.7 km/s to
allow for operations out of the 1 -sol orb it
that was established on MBC -1.

Sorties will most likely target landing sites
that had been selected ahead of time to
ensure not only science value, but also
maximize the odds of being able to land
safely due to favorable grade of the terrain,
absence of obstacles, etc. The MADV must
therefore have a reasonable ability to do a
controlled, lifting entry to adjust cross range
traject ory on the fly. A mid -L/D profile is
therefore preferred over a simpler blunt
body for this study. Lifting entry proceeds
with course corrections as needed until the

Figure 8 - WDV Concept

vehicle reaches stall speed, which is assumed
to be at or around local Mach 2 6, at which
time the MADV initiates a pitch over to a
supersonic retro propulsion (SRP) attitude
and performs a powered gravity turn and to
arrest the remaining velocity. The MADV
deploys mechanical landing gear and
touches down vertically.

The requirement for full reusability is
important to note. Inflatable aerodynamic
decelerators or other deployable
deceleration devices would result in some
savings in propellant or TPS mass, but have
no viable way to be repacked after use or
refurbished i n-situ. Parachutes could
potentially be replaced on -orbit, but would
need to be prohibitively large to arrest the
descent of a 100+ MT vehicle such as the
MADV. A fully reusable lander therefore
must rely exclusively on retro propulsion to
arrest the final velocity during descent.

4.1.2 Surface Operations
Once landed, the MADV is home for the crew
of 4 for the next 10 days of nominal
operations. For the purpose of this analysis,
we assume 50% contingency margin, or 15
days of total operational capability . 2500 kg
are allocated for equipment to support
extravehicular activities, and science
operations on the surface. MADV
ingress/egress is provided via a 2 -person
airlock, with an accompanying mechanical
lift on the exterior of the vehicle to transport
2 cr ew at a time to the surface. Consumables
are sized to allow multiple EVAs per day.
Two, 2 -person EVAs per Sol are assumed for
the purposes of this analysis.

Electrical power is generated throughout the
sortie by utilizing the boil -off from the
remaining LOX/LH2. Communications
include a low data rate connection directly to
Earth via the DSN, and high -data rate
connections to MBC via Mars -orbiting relay
satelli tes and/or directly to MBC when it
passes overhead.

4.1.3 Launch and Return to MBC
Once the sortie is complete, the MADV
becomes a single -stage -to -orbit launch
vehicle. The same LOX / LH2 engines that
were used for descent are used for ascent as
well.

The resulting total ∆V budgets and related
propellant mass fractions for the MADV are
shown below. Here again we include the
budget for the 1 -sol (design reference) orbit
as well as a 500 km circular orbit. Also note
that the numbers presented here includ e
budget for RCS maneuvers and phasing
burns to take the MADV to/from the optional
fuel resupply modules:

Table 1. Total ∆V Budgets for MADV Sorties
Return
to MBC
in:
EDL
∆V:
(m/s)
Ascent
∆V:
(m/s)
Total
∆V:
(m/s)
Propellant
Mass
Fraction
500 km
Circular
1270 4200 5470 71%
1-Sol 780 5220 6000 74%

Propellant mass fractions are calculated
using the ideal rocket equation assuming a
450 second I sp for LOX/LH2.

To put this mass fraction in perspective, the
fully loaded Space Shuttle Orbiter plus
External Tank at lift -off had a propellant
mass fraction of ~ 85%. This is an
encouraging result – even in the design -
limiting case where MBC remains in the 1 -sol
orbit, the MADV need not achieve even the
same level of dry mass efficiency as the
Space Shuttle.

5. MADV Architecture

In order to execute the mission outlined in
section IV, The MADV must be:
1) A 3 -axis controlled spacecraft capable of
both autonomous and pilot ed
rendezvous and docking operations.
2) A mid -L/D entry vehicle with high landing
accuracy
3) A vertical landing vehicle with supersonic
retro propulsion capability.
4) A single -stage to orbit launch vehicle.

It must also be fully reusable. Any hardware
that is required to be removed and replaced
between sorties must be reasonably
swapped by the MBC crew using hardware
that can be carried on MBC. (For example –
replacement of filters in the air revitalization
system between sorties is reasonable.
Replacement of a n ablative heat shield
between sorties is not). The architecture
outlined is driven in large part by this
mindset of maximum reusability and
minimum refurbishment between sorties.

5.1 Structure and Thermal Protection
System
The SLS Block IB is the assumed lift vehicle for
initial launch of the MADV. In order to
maximize available volume, the MADV is not
envisioned to be encapsulated within a
payload fairing, but will instead sit exposed
atop the SLS similar to other current rocket -
launched lifting body con cepts. It is therefore
designed to have a maximum radius at the
base of 10 m to be compatible with pla nned
SLS upper stage interfaces 7. Structures must
therefore be sized to take aerodynamic loads
and aero heating from SLS ascent in addition
to those encou ntered during Mars
operations as a result.

The Outer Mold Line (OML) of the MADV was
chosen to strike a balance between several
driving parameters:
 Must be able to perform repeated entry
operations with the same TPS materials
(no ablative coatings, etc.) So, sharp,
wing -like edges can only be utilized if
they do not create localized peak
stagnation heat rates that require single -
use TPS materials.
 Landing accuracy is considered here to be
a tradeable parameter, but landing
accuracy is desired to be as larg e as
reasonably achievable.
 Volumetric efficiency.
This trade space resulted in a mid -range
L/D lifting body. To ensure highest
achievable L/D, the MADV enters at a fairly
low angle of attack instead of as a blunt,
capsule -like body.

Initial analyses of a vehicle of this
configuration entering Mars atmosphere at
velocities between 3 and 5 m/s show that a
reusable thermal protection system is viable.
The vehicle nose and other leading edges
where peak stagnation heat rates are
greatest will likely require a carbon -carbon
composite lining, but the bulk acreage of the
entry silhouette should be able to utilize
metallic alloy skins. Optimization of entry
corridor and vehicle OML to minimize
required TPS mass is an ongoing area of
current work.

5.2 Propulsion and Power Systems
The MADV propulsion system will run
exclusively on LOX/LH2 for several reasons:
1) Highest specific impulse (~450s) of any
currently available, high -thrust chemical
rocket technology.
2) Propellants can be generated from
water with no additional chemical
reactants. This is critical to enable:

a) Minimized cost of systems to
deliver propellants from Earth
(suppliers need only deliver large
quantities of water to Mars orbit).
b) Opens the door to in -situ
production of propellants on any
celesti al body where water ice
and/or hydrated minerals are
present.

Both Main and RCS engines utiliz e Oxygen
and Hydrogen. Gaseous oxygen and
hydrogen RCS thrusters are chosen over
more traditional hypergolic -fueled thrusters
to allow RCS to be resupplied via the same
water -based supply chain that is used for the
main engines. This is critical to enable the
long -term reusability and affordability of the
MBC -based exploration effort: there exists
no viable way to generate hypergolic fuels
via ISRU, and the long -term viability of
resupply chain for MBC is assumed to be
greatest when the number and types of
resupply consumables is minimized.

The main engines are used for large -scale in -
space translational maneuvers, SRP during
EDL, and launch from the Martian surf ace.
For this study, 6 RL -10 equivalent throttle -
able LOX/LH2 main engines are assumed:

The MADV systems run off of external power
when mated to the SEP Stage, MBC or,
eventually, a refuelling module. During sortie
operations, the hydrogen and oxygen that
will naturally boil off from the propellant
tanks present a large supply of potential
energy that would otherwise be wasted.
Instead, this waste gas is used to generate
electrical power for th e vehicle.
Similar to the MBC Cryo genic Propulsion
Stage, and “Integrated Vehicle Fluids” (IVF)
LOX/LH2 -based system combines power
production, attitude control, and propellant
thermal management 8. The key tenets of this
system are:
 Waste gasses from the L OX/LH2 tanks are
burned in a hydrogen/oxygen internal
combustion engine (ICE) which drives an
electric generator to supply vehicle
electrical power.
 LOX and LH2 are pulled from the main
tanks via electric pumps and run through
a set of vaporizers and accum ulators that
continuously replenish a set of high -
pressure gaseous O2 and H2 tanks.
Thermal energy for the vaporizers is
provided from the ICE via an oil coolant
loop. This high pressure gas is then
utilized for tank pressurization, purges,
pneumatic power , and in the H 2/O 2 RCS
thrusters.

Orbital demonstrations of this integrated
system are encouraged in the near term to
ensure a system with a TRL that can support
crewed flights in the MBC timeframe. A
battery bank is also provided to ensure
adequate ener gy to keep the vehicle systems
running in the event of a temporary
generator failure or during rendezvous and
docking operations. However, fuel cells can
also be run from MADV propellant waste
gasses and remain under consideration as a
fall -back technology for power generation.
To reduce development costs and maximize
interchangeability of parts between
elements of the MBC/MADV system of
systems, all batteries and power
distribution/control components are
assumed to reuse and/or be derived from
designs curr ently in use on the Orion
spacecraft, and MADV propulsion, RCS and
power generation systems are derived from
the MBC Cryo Propulsion Stage.

5.3 ECLS and Crew Systems
The MADV E nvironmental Control and Life
Su pport system provides functions as
follows.
 Resupply of Habitable Atmosphere
(makeup gasses)

O2 makeup gasses are presumed to come
from the IVF system and makeup
atmospheric N2 is supplied via replaceable
canisters that are carried onboard MBC and
replenished between sorties.

5.3.1 Air Revitalization System (ARS):
Removal of CO 2 and Trace Contaminants
In order to maintain a breathable
atmosphere, carbon dioxide and trace
contaminants produced by the crew must be
filtered and removed. MADV ARS systems
are intended to re -use or be evolved from
systems used on the Orion – again enabling
minimized non -recur ring engineering cost
and maximum interchangeability between
Orion and MADV systems in flight.

5.3.2 Consumable Water
The crew will need consumable water for
drinking, washing, and food preparation
during the sortie. Because of the relatively
short durat ion of the sortie missions
(nominally 10 days), clothes washing and full
shower facilities are not assumed.
Figure 9 - MADV Layout

The power generation system generates
over 2 MT of byproduct water during the
sortie. This byproduct is captured and
supplies all needed consumable water for
the crew throughout the sortie. This not only
creates significant mass savings since
additional consumable water is not carried
through the vario us propulsive maneuvers,
but also reduces the complexity, mass, and
cost of the ECLS.

5.3.3 Crew Accommodations
Mass estimates include a little over 1MT for
the equipment needed to support the day -
to -day health of the crew throughout the
sortie including

5.4 Physical Configuration
Overall physical layout of the MADV is shown
in Error! Reference source not found. and
Figure 10 .

The habitable volume consists of three
decks: a flight deck, a mid -de ck with crew












Figure 10 - MADV Pressurized volume with Mid -Deck details

quarters, and an aft deck that contains the
galley and lab facilities as well the airlock for
surface access. Seats, displays and controls,
avionics, and physical layout on the flight
deck leverage common designs with the
Orion capsule to the greatest extent
possible.

The forward end of the MADV must be
covered in carbon -carbon composite for re -
entry operations, but is also the docking
interface during orbital operations. The
docking interface is therefore covered by an
articulated, TPS -covered cone. The leeward
side of the TPS below the crew cabin is also
articulated to expose radiators that are used
to help regulate avionics and crew cabin
temperatures during orbital operations.

Crew gain access to the surface through the
airlock o n the aft deck and a lift that runs
along the leeward side of the vehicle from
the airlock to the ground. Retractable
landing gear are deployed from the four aft
corners of the vehicle. The center section of
the aft end, located between the six main
engine s, is a retractable equipment lift that
lowers to give the crew access to the rovers
and other equipment for surface science
operations.

6. Conclusion

While the first Mars Base Camp mission
would only see crew operate in Mars orbit,
MBC can naturally ev olve into a true “base
camp” from which crewed missions to the
surface of Mars are launched. Via the
addition of a reusable lander / single -stage -
to -orbit launch vehicle and a fuel supply
module to support it, MBC can form the core
of an architecture that enables a vision of
humanity’s first steps on Mars. Instead of
having to commit crew to a long -duration
stay at a single location on Mars, surface
sorties from MBC are flexible, with the ability
to visit multiple surface sites in a single, year -
long missio n, and provide multiple
opportunities for the crew to return early to
the orbiting base camp in the event of a
problem on the ground.

The short -duration, scientific research
sorties in MBC -S can gradually evolve into
longer duration sorties for the const ruction
of more permanent settlements. MBC can
therefore be seen as the core system that
pivots humanity into a viable, sustainable
long -term Mars exploration program.
Testing and proving the technologies and
processes for telerobotic operations,
including sample return and analysis, at the
Moon using the D eep Space Gateway is
critical to the success of the MBC scientific
missions.

The knowledge in how to incorporate
telepresence and multiple modes of
operations into a rover as well as the
astronaut’s wo rkstation will carry forward
into the Mars robotic missions. The Gateway
will also demonstrate critical technologies,
such as solar electric propulsion. Key focus
areas for further development and research
to enable the architecture include
demonstration o f on -orbit LOX/ LH2
propellant generation, storage, and transfer,
and the characterization of potential ISRU
sources.

Figure 11 : Mars Base Camp

References
1T. Cichan, et al, Martian Moon Human
Exploration Architecture, IAC -
16.A5.2.10x35709, 67 th International
Astronautical Congress, Guadalajara,
Mexico, 2016 September.
2Cichan, T., Pratt, W., Coderre, K.,
“International, Scientific, and Commercial
Opportunities Enabled by a Deep Space
Gateway,” IAC, 2017
3Hopkins, J.B., “Early Telerobotic
Exploration of the Lunar Farside Using Orion
Spacecraft at Earth -Moon L2,” GLEX -
2012.02.3.2x12595, Global Space
Exploration Conference (GLEX) 2012,
Washington, D.C., 2012, 22 – 24, May.
4NRC -2007, National Research Council ,
2007, The Scientific Context for Exploration
of the Moon, National Academy Press,
Washington, DC, pp59 -60.
5Burns, J., “Science from the Moon: The
NASA/NLSI Lunar University Network for
Astrophysics Research (LUNAR)”, White
Paper submitted to the Planeta ry Sciences
Decadal Survey.
6Steinfeldt, B.A., Theisinger, J.E., Korzun,
A.M., Clark, I.G., Grant, M.J., Braun, R.D.,
“High Mass Mars Entry, Descent, and Landing
Architecture Assessment,” AIAA -2009 -6684.
7NASA., “Space Launch System,” Fact
Sheet FS -2016 -02 -04 -MSFC, 2016, George C.
Marshall Space Flight Center,
https://www.nasa.gov/marshall
8Zegler, F.., “An Integrated Vehicle
Propulsion and Power System for Long
Duration Cryogenic Spaceflight,” AIAA, 2011
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