Lunar Lander Concepts for Human Exploration

Abstract

A new generation of lunar lander is to be the reference payload for the NASA Ares-V heavy-lift launch vehicle, still in conceptual development. The surface-payload capability of the lander is primarily a function of propulsion choice, staging method, and configuration choice. A variety of staging methodologies are investigated, and the benefits and disadvantages of staging in low lunar orbit and staging later in the final descent burn are presented, as are the benefits of dropping tankage before touchdown to reduce the lander size and mass. Storable and methane propellents for the ascent burn are evaluated. A variety of configuration options are presented, and the discussion includes the context for downloading heavy payloads for outpost buildup. The transportation architecture variations assume the basic NASA Exploration Systems Architecture Study architecture, and the surface operations are traded to match compatible lander configurations.

Full text

Lunar Lander Concepts for Human Exploration
Benjamin B. Donahue,∗Glenn N. Caplin,†David B. Smith,‡John Behrens,§and Curtis Maulsby¶
Boeing Advanced Systems, Huntsville, Alabama 35806
DOI: 10.2514/1.29270
A new generation of lunar lander is to be the reference payload for the NASA Ares-V heavy-lift launch vehicle, still
in conceptual development. The surface-payload capability of the lander is primarily a function of propulsion choice,
staging method, and configuration choice. A variety of staging methodologies are investigated, and the benefits and
disadvantages of staging in low lunar orbit and staging later in the final descent burn are presented, as are the benefits
of dropping tankage before touchdown to reduce the lander size and mass. Storable and methane propellants for the
ascent burn are evaluated. A variety of configuration options are presented, and the discussion includes the context
for downloading heavy payloads for outpost buildup. The transportation architecture variations assume the basic
NASA Exploration Systems Architecture Study architecture, and the surface operations are traded to match
compatible lander configurations.
Introduction
THE lunar lander occupies a unique position within NASA’s
Constellation programs: it is the interface between the two
primary system architectures: the transportation and surface systems.
And in some scenarios, the lander performs functions for both
architectures. Subsequently, the design of the lander must satisfy the
requirements established by the two architectures. Optimization of
the design not only satisfies the requirements, but alsomaximizes a set
of criteria (i.e., figures of merit). Alternative designs are established
by examining and evaluating differentpotential designs within a trade
space. Although the trade space is broad, resulting in numerous
potential designs, this paper explores three critical portions of the
trade space: the staging approach, ascent-propulsion technology, and
general arrangement. Although there are numerous potential figures
of merit, the discussion here is limited to evaluation of one of the
primary performance parameters as part of the transportation
architecture: payload mass delivered to the lunar surface.
Translunar Injection and Lunar Orbit Insertion
A new generation of lunar lander is to be the reference payload for
the NASA Ares-V heavy-lift launch vehicle, still in conceptual
development. After solid rocket booster drop-off and core stage
separation, the Ares-V’s second stage finalizes the burn to place the
lunar lander into a 160 by 160 nm, 28.5-deg, low Earth orbit (LEO).
Subsequently, the crew exploration vehicle (CEV), placed in orbit
with the Ares-I, docks with the lander. The Ares-V’s second stage,
still attached to the lander, serves as the Earth-departure stage (EDS).
From LEO, with 40% of its propellant load remaining, the EDS will
inject both the lander and CEV into a three-day transfer. Later, the
crew will depart back to Earth usingthe CEV’s service-module (SM)
propulsion system. An illustration of the Ares-V is given in Fig. 1. In
Apollo, the command-module/SM was launched together with the
lunar excursion module (LEM) on a single Saturn-V. The Saturn’s
suborbital-start third stage also served as the EDS. Like Apollo, the
Constellation CEV SM uses storable propellant for trans-Earth
injection (TEI), but unlike Apollo, it will not provide propulsion for
the lunar orbit insertion (LOI) burn; this is instead done by the lander,
with LOI propellant contained in the descent tanks. Offloading the
LOI propellant from the SM keeps the CEV from exceeding the lift
capability of Ares-I. The CEV is projected to weigh 20 t in LEO. The
Ares-V second stage/EDS, weighing about 247 t fully loaded,
expends about 60% of its propellantboosting the lander to LEO. Once
in orbit (and after the docking of the CEV), the second stage fires
again to boost the 65-t lander/CEV stack to translunar injection (TLI).
Descent and Ascent
After lunar orbit capture, the lander separates from the CEV and
initiates descent from a 100-km circular low lunar orbit (LLO) into a
100 by 15 km elliptical phasing orbit; at 15 km, initiation for final
descent begins. By the time the lander nears touchdown, all of its LOI
and its descent propellant has been burned off and the vehicle is
significantly less massive than when it started its LOI burn. Because
of this, significant engine throttling is required. (NASA Marshall
Space Flight Center is presently investigating deep-throttling
technology, including RL-10-derived and Pintle injector engine
options.) Following surface operations, the crew boards the ascent
stage, which ascends to LLO and docks with the CEV. In Fig. 2, the
NASA Exploration Systems Architecture Study (ESAS) lander is
pictured. Mission delta-velocity (delta-V) values used in this paper
are given in Table 1. Note that this paper is focused on lander designs
that support the outpost build, and the LOI delta-V for the polar
mission is used.
Several other top-level requirements are imposed on the design
beyond the preceding delta-V and mission operations: specifically, a
shroud diameter of 10.0 m maximum, a low-impact docking system,
four crew members for lunar missions, a 100-kg minimum payload
return from the lunar surface to Earth, and a surface air lock.
In addition to its role in delivering crew and cargo to the lunar
surface, it is desirable to have a variant of the basic lander serve in a
cargo-only mission role. For these missions, only the Ares-V launch
is required, which places up to 53.6 t to TLI. The remainder of the
mission is as previously described, but without the crew or ascent
vehicle.
Surface Architecture and Requirements
Although the transportation architecture and requirements, as
previously described, are fairly well-established, the surface
architecture, and hence the requirements for the lunar lander as the
interface to the surface architecture (and possibly an element of the
surface architecture), is less mature. Surface architectures range from
the following:
Presented as Paper 7443 at Space 2006, San Jose, CA, 17–21 September
2006; received 12 December 2006; revision received 11 October 2007;
accepted for publication 11 October 2007. Copyright © 2007 by The Boeing
Company. Published by the American Institute of Aeronautics and
Astronautics, Inc., with permission. Copies of this paper may be made for
personal or internal use, on condition that the copier pay the $10.00 per-copy
fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers,
MA 01923; include the code $10.00 in correspondence with the CCC.
∗Propulsion Engineer, Flight Engineering, Advanced Systems, MC JV-05,
950 Explorer Boulevard. Senior Fellow AIAA.
†Systems Engineer, Space and Intelligence Systems, MC W-S10-S356, El
Segundo, CA 90245
‡Manager, Space Exploration Systems, MC 793C-G030, Arlington, VA.
§Structural Engineer, Space and Intelligence Systems, MC W-S10-S356,
El Segundo, CA 90245
¶Mechanical Design, Advanced Systems, MC JV-05, 950 Explorer
Boulevard.
JOURNAL OF SPACECRAFT AND ROCKETS
1

Surface architectures range from those in which the lander plays a
relatively passive role to architectures in which the lander is highly
active. The former is a surface system infrastructure in which cranes
and trucks offload, transport, and assemble the lunar outpost from
cargo delivered to the surface. This category might also include
landers that provide a relatively minor assist to the surface
architecture, such as provide partial offloading of cargo, but that still
rely on a surface infrastructure and play no other role. At the other
end of the spectrum, the high active role, the lander may include
mobility and docking features to provide surface transportation and
assembly capability, as well as functionality with the constructed
lunar outpost (i.e., reuse of the lander in some capacity such as
descent tanks providing oxygen storage). There are roles that are
between passive and highly active, such as the lander serving as part
of a loosely connected lunar colony without surface transportation
capability.
The lander can be further influenced by the surface architecture
plans for outpost build regarding the use of dedicated cargo missions
versus mixed crew/cargo missions (for example, if the build
architecture preemplaces a habitat module close to the surface,
surface access from the lander on subsequent crewed missions would
be less of a concern). The lander designs described make no
assumption regarding surface architecture or impose architecture-
derived requirements, but some designs will be noted as capable of
supporting particular surface architectures.
Lander Surface-Payload-Delivery Capability
As already noted, the figure of merit considered for this paper is the
payload mass delivered to the surface. The lander’s payload-delivery
capability is a function of launch vehicle capability, propulsion
efficiency, staging method, and lander mass exclusive of the payload
(i.e., structure, mission requirements, and mission design). For
purposes of this trade study, the mission requirements and design are
assumed to be fixed from the ESAS report. Finally, the performance
for this paper is assumed to be independent of lander general-
configuration trades. Although the particular lander configuration
will, of course, impact mass, the evaluation regarding propulsion
technology and staging is simplified with this assumption.
Lander Engine Technology (Propellant-Type) Trades
In Fig. 3, information is listed for the Ares-V upper-stage engine
(RS68), The lunar lander descent-stage engine (RL10-B2) and a
potential-ascent engine (Bell 8258 LEM ascent engine). Two
propellant types are considered in the trade for ascent: storable
nitrogen tetroxide (N2O4) with monomethyl hydrazine (MMH) and
oxygen (LO2) with methane (CH4). No trade is performed regarding
the LOI and descent propulsion; LO2with hydrogen (LH2)is
baselined. The engine Isp, nozzle area ratio , and engine thrust-to-
weight ratio T=W are shown in Table 2.
The RL10-A4 is a production O2=H2engine; its demonstrated Isp
is 451 s, but for this analysis, the Isp is reduced by 5 to 446 s, for
conservatism. For the ascent’s all-cryogenic O2=CH4option,
advanced passive insulation, including vapor-cooled shields and
multilayer insulation (MLI) is used to limit boil-off on the surface.
Additional cryocooler units may be required, specially should a
breach in the integrity of the passive insulation occur; these were
added in the calculations to provide additional conservatism (later
missions may feature south pole stays of up to six months). Recent
NASA Johnson Space Center analysis indicates the capability of
O2=CH4systems to use the shared main and reaction control system
(RCS) propellant tanks.
Lander Staging-Methodology Trades
The second major element of the trade space to consider is the
staging strategy. In principle, this could range from a single stage to
perform all propulsive events to a separate stage for each propulsive
event (and variants thereof, such as drop-tank stages in which only
the tanks are jettisoned, sometimes referred to as a half-stage, or
propulsion systems that are used for certain events but are not staged
away). Table 3 illustrates a trade space for staging. Each of the major
propulsion events is considered with the potential staging
combinations. Note that the table divides the braking burn into two
phases: the initial 1693-m/s-descent braking burn that removes most
of the velocity and a 300-m/s final-descent-to-touchdown burn.
Fig. 1 Ares-V launch vehicle.
Table 1 Lander delta velocity
Burn Name Delta-V, m/s Stage Burn Name Delta-V, m/s Stage
1 Earth departure 3327 EDS 4 Descent 1963 Descent
2 Outbound midcourse 20 Descent 5 Ascent 1905 Ascent
3 Lunar orbit insertion 892 Descent 6 Trim and rendezvous 60 Ascent
Fig. 2 NASA ESAS report lunar lander concept.
Fig. 3 Reference Ares-V upper-stage, lander descent-stage, and
ascent-stage engines.
2DONAHUE ET AL.

Table 3 indicates, for each major propulsive event, which stage
performs the event. LOI values of 1.5, 2.5, and 3.5 indicate that the
event is performed by a drop-tank stage with the tanks present, and
the whole number for the following event indicates that the tanks
were dropped. For example, 2.5 for LOI followed by 2 for descent
indicates that tanks were dropped following the LOI burn; the full
complement of tanks was present for LOI, but only a partial set of
tanks are used for the descent: that is, the tanks were dropped
following LOI. By inspection, a number of possibilities can be
discarded, at least for an initial evaluation; single-stage versions
have, in the past, provided insufficient performance (without surface
refueling) and have no independent abort capability (i.e., ascent
punchout). Providing a separate stage for the final descent (last
300 m=s) is probably not a good mass trade, although doing so
preserves the ascent stage for ascent-only use if that becomes
required. This leaves the 2- and 2.5-stage configurations (4 of the 13
from the trade table) as a reasonable starting point for staging trades
and is discussed in the remainder of the paper.
Two-Stage Surface Staging
The two-stage surface-staging concept is the classic Apollo
approach, which is also the current nominal ESAS report design and
reference design for this paper. The system is referred to as surface
staging, because that is where the staging event occurs (upon
initiation of the ascent burn). We will not discuss this approach in
detail, because it is well-documented, but note a few advantages:
1) The independent ascent stage provides an abort capability
independent of the descent propulsion system.
2) Ascent propulsion only requires a single start (i.e., no prior use
of the ascent system before ascent).
3) There is only one staging event, which occurs in a benign
environment.
4) There is no dropped debris to control impact.
The primary disadvantage relates to the large tank volume
required to support both the LOI and descent burn. This presents
configurational challenges relating to surface access and cargo
offloading. These may be nonissues, depending upon the surface
build architecture, particularly if it is based on a construction
infrastructure (i.e., cranes available to offload the payload). In
addition, a substantial throttle range is required for the descent
engines, due to the large change in lander mass as LOI and descent
propellants are expelled.
Drop-Tank-Staging Concepts
Drop-tank-staging variants attempt to ameliorate some of the
issues associated with the surface-staged designs: specifically, the
configurational advantage of staging a substantial portion of the tank
volume before landing. The tanks may be dropped either following
the LOI burn (in LLO) or during descent (with the lander final-
descent tanks serving as feeder tanks from the drop tanks to avoid a
descent-engine restart). In either scenario, the dropped-debris
scenario remains a disadvantage. The concepts also seek to gain
some payload mass advantage that comes from staging away empty
tankage.
Descent-Staging Concepts
The descent-staging option consists of a separate, complete,
dedicated stage that is used for the LOI burn and most of the descent
burn. This common LOI-descent stage (CLDS) would be roughly the
size of a Delta-IV upper stage. Once staged 4 km above the surface,
the terminal descent–ascent stage descends the final 300 m=sto
touchdown. Terminal propellant is contained in the ascent tanks, and
the engine(s) used for the final descent may be reused for the ascent.
This approach reduces the physical size of the landing craft, when
compared with the reference ESAS lander, and significantly
increases its surface-payload capability.
There are other advantages, including a much-reduced throttling
requirement for final descent. The Surveyor robotic lunar lander used
descent staging in the 1960s. In Fig. 4, a plot of terminal descent
altitude vs range to go is given. The lander is approximately 2 km
above the empty CLDS at the time of its impact; the lander touches
down about 4 km downrange of the impact point. Both distances are
subject to trade; increasing the separation requires either earlier
separation or a shallower flight path, either of which requires
additional descent propellant, which adversely affects the delivered
payload. In Fig. 5, the CLDS lander separation maneuver is shown.
The advantages of descent staging are many. First and foremost,
descent staging removes most of the LOI and descent burn tankage
before touchdown. This provides for a much smaller lander,
significantly increases lander payload capability, and simplifies
surface access. It also allows for similar vehicle T=W requirements
for final descent and initial ascent, allowing common engines. The
terminal descent stage requires some throttling, though not as
significant as the deep-throttling requirement of the reference ESAS
mission lander. Descent staging allows the separate CLDS to use a
Table 2 Lunar lander engine parameters
Name Type Isp T=W Feed system Notes Use
1N2O4=MMH Storable 329 200 35.0 Pressure Hypergolic Ascent
2LO2=LCH4Light cryogen 360 200 35.0 Pressure New design Ascent
3LO2=LH2Deep cryogen 446 84 60.3 Pump RL10-A4 LOI/descent
Table 3 Lander staging trade space
Burn LOI Descent braking burn Descent final Ascent Comments
Delta-V, m/s 892 1663 300 1905
1 stage 1 1 1 1 Single stage does all
1.5 1 1 1 Drop tanks after LOI
1.5 1.5 1 1 Drop tanks after braking burn (at 4 km)
1.5 stage 1.5 1.5 1.5 1 Drop tanks after landing (on surface)
2111
2 2 1 1 CLDS design
2 stage 2 2 2 1 ESAS/Apollo
2.5 2 2 1 Drop tanks after LOI
2.5 stage 2.5 2.5 2 1 Drop tanks after most of the descent burn
3 2 2 1 Approximately same size LOI/descent stages
3 stage 3 3 2 1 Only 300 m=sfor stage 2
3.5 stage 335 3 2 1 Drop tanks after LOI, only 300 m=sfor stage
2
4 stage 4 3 2 1 Only 300 m=sfor stage 2
DONAHUE ET AL. 3

long-nozzle, high-Isp (458 s) engine, and the engine for this stage
does not require any throttling. This separate CLDS would be similar
in size to an existing upper stage, allowing it to come off a common
assembly line.
There are disadvantages to descent staging. First, it requires a
time-critical engine start of the terminal stage 4 km above the surface,
and the engine must be restarted for ascent. Also, descent staging
leaves a spent stage on the surface 4 km away from the landing site.
Performance Assessment
Performance (payload mass to the surface) was assessed as a
function of the staging concept and ascent propellant. A common
ascent payload is used in all cases, as described next. The two-stage
surface-staged lander is used as a reference. The mass of the ascent
cab for a crew of four is determined from a series of interdependent
algorithms that capture the intricacies of habitat subsystems as
functions of the number of crew, duration, volume, radiation
shielding, internal pressure, redundancy, spares, and other
considerations. Additional masses of 400, 440, and 125 kg are
allocated for docking port, crew/effects, and extravehicular activity
(EVA) suit masses, respectively; total ascent-crew-module payload
is 2655 kg. Reasonable propellant margins, residuals, tank gage
uncertainty, boil-off, dry-weight growth, gravity losses, throttling
losses, and other margins are applied. For this analysis, only 2- or 2.5-
stage landers were considered; subsequent briefings may address
single-stage and stage-and-a-half concepts.
Areference crew-mission lander uses a single O2=H2RL10-A4
engine for LOI and descent and a single pressure-fed N2O4=MMH
engine for ascent. Its total mass is 45 t (the maximum that the Ares-V
EDS can boost to TLI), and it delivers, in addition to the ascent stage,
a surface payload of 4.8 t. The lander performance model optimizes
the surface-payload mass within the boundaries defined by the
assumptions and the various propulsion and staging options. The
surface payload is the variable to be maximized given fixed CEV and
Ares-V to TLI mass values (see Table 4). From this reference case,
ascent-propellant and staging-mode trades were run. Ascent stages
use a single pressure-feed engine (either storable or O2=CH4) and
ascend to a 100-km LLO.
For a cargo-only case, in which no CEV or ascent stage is carried,
the reference ESAS descent stage, if unchanged, could carry 16.4 t of
surface cargo (Table 4). This descent stage is the same as the crew-
mission descent stage (the LOI/descent propellant load is less and the
propellant split is different). However, at this capability, it does not
take full advantage of the Ares-V capability to TLI, which, if the
CEV is not taken, is 53.6 t.
A redesigned descent stage, sized to maximize the surface payload
and take full advantage of the Ares-V capability, could deliver 19.8 t
to the surface. This is a descent stage with increased thrust, larger
descent tanks, and a heavier frame than the reference crew-mission
descent stage.
Figures 6 and 7 show how crew-mission lander mass varies with
ascent-propellant choice and staging mode; the first two bars refer to
ESAS staging cases, the middle two bars refer to drop-tank-mode
cases, and the last two bars refer to descent-staging-mode cases. For
each pair, the first bar refers to the storable-propellant-ascent-stage
case and the second bar refers to the methane-ascent-stage case.
Surface payloads range from 4.8–5.0 t (ESAS), 5.2–5.4 t (drop tank),
and 5.6–6.1 t (descent staging). Methane-ascent propulsion provides
a slight (0.2–0.4-t) increase in surface payload over N2O4=MMH
(Fig. 7). The drop-tank mode provides a payload gain of 0.4 t
compared with the ESAS reference, whereas a larger gain (0.9–1.1 t)
is achieved with descent staging. The drop-tank- and descent-
staging-mode landers are physically smaller than the reference.
Figure 8 illustrates lander propellant tank volumes for each of the
three staging modes. Only the “landed”tank volume is shown. For
each of the staging modes, two values are given: the landed descent-
stage tank volume is given first and next to it is the ascent-stage tank
volume. For the ESAS case on the left, 70 m3(cubic meters) of
LO2=LH2LOI/descent tank volume is landed, along with 4m
3of
N2O4=MMH ascent tank volume. For the drop-tank mode in the
center (with all LOI propellant in the jettisoned tank), 32 m3of
LO2=LH2tankage is landed (4m
3for N2O4=MMH ascent tankage):
less than half the volume of the reference ESAS lander volume.
Finally, for the descent-staging mode (right), the all-LO2=LH2
CLDS is not landed; only the N2O4=MMH terminal descent–ascent
Fig. 4 Descent-staging: altitude vs range to go.
Fig. 5 Descent-staging: separation maneuver (MECO is the main
engine cutoff).
Table 4 ESAS mode reference lander mass statement, 65-t ARES-V to TLI value
Element Crew mission (ref.) Cargo common descent stage Cargo optimized lander Comments
To TLI Total 65,000 45,000 53,600 Ares-V Earth-departure-stage payload
CEV In LEO 20,000 0 0 In LEO, service module with TEI propel
Lander Total 45,000 45,000 53,600 In LEO (Ares-V upper-stage payload)
Dry mass 6756 6756 7710 Single-engine, 446 Isp RL10A4 (84 ER)
Descent prop 11,855 13,196 15,735 LO2=LH2
LOI prop 12,004 8302 9899 LO2=LH2
RCS prop 385 348 416 280-s Isp, storable
Descent Payload 4781 16;350 19;840 Left on surface
Dry mass 2320 n/a n/a Single-engine, pressure-fed, 329 Isp
Ascent prop 4133 NTO/MMH, storable
RCS prop 111 280-s Isp, storable
Ascent Crew cabin 2655 Cab, port, crew, EVA suits, etc.
4DONAHUE ET AL.

stage is landed and its total tank volume is 9m
3: an order-of-
magnitude reduction in landed propellant tank volume compared
with the reference. The benefit is a much smaller, shorter, and lighter
lander that does not require the deep throttling that is characteristic of
the larger ESAS lander.
Sensitivity to Ares-V EDS Capability
For Ares-V EDS mass to TLI values of 60, 62.5, and 65 t, lander
total mass values are 40, 42.5, and 45 t. In Table 5, lander payload
mass is given. Ascent mass values are 9.2 and 9.0 t for storable and
LO2=methane propellants, respectively. Delivered descent-stage
surface payloads (in addition to the ascent stage) are 2.81–3.03 t for
the 40-t lander, 3.79–4.02 t for the 42.5-t lander, and 4.78–5.01 t for
the reference 45-t lander. Values for the other two staging modes are
also listed.
Lander Configuration Options
The lander configuration trades examine the arrangement of
equipment that achieves the lander requirements. These general
arrangements are dominated by the larger equipment: notably, the
ascent module, tanks, engines, and payload. Requirements are first
dictated by the transportation and surface architectures, then further
derived in terms of particulars such as tank size, staging, etc., as
described in the previous section. Because these trades are still open
and the surface architecture is being refined, the configuration option
set is quite large and a survey of potential designs is presented.
Typically, the trades involve evaluating one figure of merit against
another (e.g., delivered mass vs accessibility).
As discussed earlier, the role of the lander in the surface
architecture can influence its design:
1) The passive lander has the salient design characteristic of
readily offloadable payload (i.e., good clearances for access and
removal of payloads). In some cases, this may be aided by
mechanisms on the lander.
2) The active lander has the salient design characteristics of
payload (such as habitat modules) close to the surface, provisions for
self-powered surface transportation, preferably smaller to aid in
access and transportation issues, and adequate engine clearance for
surface transportation.
3) The colony lander has the salient design characteristic of
payload (such as habitat modules) close to the surface.
For each staging strategy, configurations can be developed and
evaluated in regard to how well they will tend to support a particular
surface strategy (ascent-propellant trades for those propellants
discussed have a relatively minor impact on the configuration trades).
An approach to evaluating and shaping the lander design is shown
in Fig. 9. The figure illustrates a trade tree that describes various
options in regard to creating a lunar outpost and, ultimately, the role
of the lander in the creation of the outpost (beyond the role of
transporting the crew and cargo from LLO to the surface). As shown
in the figure, the highest level of trade is the basic architecture for the
lunar outpost. Two basic alternatives are suggested, although there
are, no doubt, others. The term “tightly coupled”refers to a lunar
outpost built from modular components, which are subsequently
connected together on the lunar surface through a hard connection,
and likely includes a pressurized interconnect between at least some
of the modules. The loosely coupled surface architecture either has
no connection among the surface elements or relatively little
connection (such as power). In other words, each lander provides a
relatively self-sufficient habitat, and so the collection is a “colony,”
which approximately suggests that the landers are relatively close to
each other, some small number of meters implying that they were
0
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(CLDS)
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Fig. 6 Lander masses vs staging and ascent-propellant choices; 45-t lander.
ESAS Ref
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Desc Pa
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load 4,780 5,010 5,150 5,380 5,600 6,050
ESAS ESAS Drop Tank Drop Tank Desc Stged Desc Stged
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Drop Tank
Staging
Descent
Staging
Descent Surface Payloads (Excludes Ascent Stage)
0
1000
2000
3000
4000
5000
6000
7000
Descent Payloads Mass, (kg)
Desc Pa
y
load 4,780 5,010 5,150 5,380 5,600 6,050
ESAS ESAS Drop Tank Drop Tank Desc Stged Desc Stged
NTO/MMH
asc
O2/CH4
asc
.. NTO/MMH
asc
O2/CH4
asc
.. NTO/MMH
asc
O2/CH4
asc
Fig. 7 Surface payload vs staging and ascent propellant; 45-t lander.
Fig. 8 Landed propellant tank volume comparison; 45-t lander.
DONAHUE ET AL. 5

transported from the landing site. Scattered colonies have no surface
transportation; the outpost simply consists of a number of landers
that remain in their original landing site, the distance between them
determined by the minimum safe landing separation. The tightly
coupled architecture is usually what is associated with lunar outposts,
although a loosely coupled architecture is not precluded.
The next level of the trade tree evaluates the construction method:
either via emplacement of basic construction infrastructure (cranes or
other offloading equipment and surface transportation) or no
infrastructure is required, in other words, the lander provides surface
mobility, leveling, connectivity to another node of the outpost, etc.
Self-transportation eliminates the need for designing and emplacing
an infrastructure and allows outposts built with the fewest number of
flights, but also requires that each lander (or payload) provide its own
mobility features. The mass for mobility features is estimated at
about 1.2 t (wheels, motors, and leveling features found on the
surface transporter). The next layer further defines the lander role in
the construction and outpost architecture (either the whole lander is
used or some cargo is offloaded). In some architectures, a surface
transporter can drive under lander surface payloads and transport
these payloads to the base and then return to service the next lander’s
payload. The final two layers of the trade relate to trades regarding the
lander itself once its role in the overall surface architecture has been
established. The staging trade discussed earlier is critical, because it
dictates the tankage at the surface (drop tanks are included in the
descent-staging category). Finally, there is an entire tree that relates
to the physical trades for the lander.
At this point, no assumptions are made regarding the trades at the
higher level of the trade space, and the discussion merely serves to
illustrate the linkage between the surface architecture and lander
design.
The trade space for the lander configuration includes 1) general
order and arrangement (i.e., stacking), primarily of the surface-
payload module, ascent module, tankage, and main engines;
2) orientation of the primary axis at landing (horizontal or vertical);
3) tankage trades (multiple tanks provide flexibility, single tanks
provide greater efficiency, particularly for cryogenic thermal control,
and toroidal tanks provide volumetric efficiency in some
configurations); 4) fixed configurations vs post launch deployed or
reassembled in orbit; and 5) habitat or other major cargo shape.
Clearly, the trade space is quite large. As an illustration of a subset,
Fig. 10 lists elemental drawings of a diverse set of lander
configurations. All sketches are representative. Five types are
illustrated and the title of each corresponds to the position of the main
payload module. The module may be placed on the top (types 1 and
3), sides (type 2), bottom (type 5), or axially integrated into the center
(type 4). Main engines may be located at the bottom center (types 1,
2, 3, 4, 5B, and 5C), sides (type 5A), ends, or corners. There may be
twin modules [types 2A, 2B, 3 (right), and 5C]. Propellant tanks may
be mounted on the sides (5A), top (5B and 5C), or bottom (types 1,
2A, 3, and 4).
Undercarriage concepts (type 5), locating their payload on the
bottom, allow ease of cargo offloading; representatives in Fig. 10 are
5A–5C. The cargo is lowered to the surface by hoists. Engine exhaust
shields are not shown. A module preintegrated with a surface
transporter may be ready for use immediately after touchdown (5A).
Outpost buildup might be simplified compared with other concepts,
because modules could be offloaded, transported, and mated together
to form a surface base without the use of cranes or ramps.
Undercarriage configurations also provide short ingress/egress
paths for personnel and a low vehicle c.g. lessens the tip-over hazard
when landing on sloped terrain. Engine-out options include opposed
engine shutdown and/or RCS assist. Missions carrying a single large
payload benefit the most from this approach. Proponents cite its
cargo positioning; opponents cite its spread-engine arrangement,
because it is not directly under the c.g. (though the sum of the thrust
component is through the c.g.). As in all designs, engines gimbal to
track the changing c.g. as propellant is consumed. Type 5B differs
from 5A in that the engines are centered, and at the bottom of the
vehicle, the ascent and descent tanks are positioned above the crew
cab (rather than to the sides, as in 5A). Type 5C locates twin modules
just outboard of bottom-center-positioned engines (the tankage is
Table 5 Descent surface-payload variation with total lander mass (all mass is in metric tons)
ESAS mode Drop-tank mode Descent-staging mode
Storable-ascent stage Methane-ascent stage Storable-ascent stage Methane-ascent stage Storable-ascent stage Methane-ascent stage
Lunar
lander
mass
Desc surf
payl
Asc-stage
total
Desc surf
payl
Asc-stage
total
Desc surf
payl
Asc-stage
total
Desc surf
payl
Asc-stage
total
Desc surf
payl
Term
desc–
asc
Desc surf
payl
Term
desc–
asc
45.0 4.78 9.22 5.01 8.99 5.15 9.22 5.38 8.99 5.60 14.4 6.05 14.0
42.5 3.79 9.22 4.02 8.99 4.14 9.22 4.37 8.99 4.75 14.0 5.18 13.6
40.0 2.81 9.22 3.03 8.99 3.12 9.22 3.35 8.99 3.91 13.6 4.32 13.2
Top-level surface
architecture
architecture
Construction
Lander
role
Staging
Tightly coupled
(Lego)
Outpost
Loosely Couple d
(Colony)
Ve r y
proximate Scattered
Self de ployed
cargo
No infrastruc ture,
se lf b uil d
Wh ole
Lander
Surface
sta g e d
Descen t
sta g ed
Surface
sta g ed
Descent
sta g e d
Similar to tightly
coupled trade tree
Lander assisted
unload
Passive Lander,
Crane unload
Surface construction
infra structure
Surface
sta g e d
Descen t
sta g ed
Surface
sta g ed
Descen t
sta g ed
Wh ole
Lander
Surface
sta g ed
Descent
sta g e d
Tightly coupled
(Lego)
Outpost
Loosely Couple d
(Colony)
Ve r y
proximate Scattered
Loosely coupled
(colony)
Very
proximate Scattered
Self de ployed
cargo
No infrastruc ture,
se lf b uil d
Wh ole
Lander
Surface
sta g e d
Descen t
sta g ed
Surface
sta g ed
Descent
sta g e d
Self-deployed
cargo
No infrastructure,
self-build
Whole
lander
Surface
sta g e d
Descen t
sta g ed
Surface
sta g e d
Descen t
sta g ed
Surface Descent Surface
sta g ed
Descent
sta g e d
Surface
sta g ed
Descent
sta g e d
Surface Descent
Lander-assisted
unload
Passive lander,
crane unload
Surface construction
infrastructure
Surface
sta g e d
Descen t
sta g ed
Surface
sta g e d
Descen t
sta g ed
Surface Descent Surface
sta g ed
Descen t
sta g ed
Surface
sta g ed
Descen t
sta g ed
Surface Descent
Whole
lander
Surface
sta g ed
Descent
sta g e d
Surface
sta g ed
Descent
sta g e d
Surface Descent
staged staged staged staged staged staged staged staged staged
staged
Fig. 9 Surface architecture trade tree.
6DONAHUE ET AL.

directly above), and type 5D (not shown) locates the engines and
tankage on the vehicle ends.
Top-loaded concepts (type 1) require large offloading cranes to lift
cargo modules off the top. If the module is heavier than the crane,
there is a danger that the module will pull over the crane, unless the
crane is so large that it can straddle the lander. The cost to launch,
emplace, transport, and service a straddler crane of this size
(physically larger than the lander) may become problematic. The cost
and complexity of a three-story-high straddler crane/offloader may
be a sizable disadvantage of this “lift off from the top”option. If
modules are not to be downloaded and only single-module outposts
are envisioned (or colony-type outposts), then type 1 might be
preferred. To lessen the difficulty of offloading payload and allowing
top placement without requiring a crane, the top-hinged concept
(type 3) may be considered.
Top-hinged concepts do not require an offloading crane, though
they require a mechanism and hinged-cradle system to rotate the
module off the top, around the sides, and onto the surface. In Fig. 10,
type 3, both single-module (left) and dual-module versions (right) are
shown.
Side-loaded concepts do not require cranes, though the cargo must
be split into two equally weighted pieces to retain symmetry and
proper c.g. for flight control during the descent.
Center-loaded concepts also do not require an offloading crane,
and the main cargo can be one piece. The type-4 central-habitat
configuration locates the module vertically and in the center of the
descent stage. The illustration shows the module surrounded by
tankage (except on one side), with main engines directly underneath,
in the center. Once landed, the module is rotated down from a hinge
point at the bottom until it is horizontal.
Fixed-habitat concepts are not designed for habitat removal. A
centrally located habitat is either surrounded by descent propellant
tanks or propellant tanks are located above. To allow for close
surface proximity, descent engines are located at the vehicle corners,
leaving the center position for the habitat.
Specific Configurations
With the requirements, trade space, and typical high-level
configurations identified, a few landers were defined to the next level
of detail. The following describes a few specific configurations over a
broad range of the trade space.
Atop-mounted configuration (Fig. 11) is similar to the ESAS
lander pictured in Fig. 2. This configuration would likely support a
tightly coupled outpost with the aid of some surface infrastructure.
An ascent stage is located on the top, with its single ascent-engine
nozzle sitting in a central void that runs through the descent stage,
which has circumferential descent tanks. The ascent stage features a
cylindrical crew cabin, with tanks positioned on the sides. The
descent stage may be common for both crew and cargo (Fig. 12)
variants.
Undercarriage concepts have the virtue of providing payload
proximate to the surface, either for simplified offloading (which may
include self-transportation) or accessible modularity for an outpost
built using the entire lander. Figures 13 and 14 illustrate two
undercarriage concepts.
In Fig. 13, an undercarriage-cargo configuration featuring a
doughnut-shaped surface habitat is shown. A single descent engine
occupies the central void in the habitat. The configuration features a
single descent hydrogen tank with several oxygen tanks located on
the side (the large single hydrogen tank is more easily insulated than
several smaller hydrogen tanks). An ascent stage sits atop the
hydrogen tank and features two side-mounted engines. This
configuration requires ascent-module EVA upon arrival and
departure; for all remaining surface activities, the crew has a short
egress path to the surface. The addition of surface mobility and
docking features (not shown) would permit this lander to support
self-built, tightly coupled, outpost architectures.
A second undercarriage configuration is illustrated in Fig. 14,
featuring side engines and descent tanks. The illustration does not
show the LOI drop tanks (only descent tankage is shown). This
concept is suited to delivery of cylindrical payloads that may initially
be preintegrated with wheeled transports. The transport drives up
under the payload, which is lowered onto the transporter and is driven
off to be jointed to other payloads. This would facilitate the buildup
of a base made from joined modules, without cranes, simplifying site
buildup. A variation of concept 2 is given in Fig. 15; pictured is a
cargo-only vehicle that features full LOI and descent tankage, no
CLDS is used.
Fig. 10 Cargo lander configuration-options montage.
DONAHUE ET AL. 7

A LOI drop-tank configuration is shown in Fig. 16. It features a
vertical, central, cylindrical surface habitat. Around it, LOI and
descent tanks are positioned. The cylindrical habitat is divided into
two sections. The bottom is doughnut-shaped, with a central void
occupied by the descent engine. The lower pressurized section
contains the surface air lock and storage areas; the section above the
central void is a pure cylinder and contains the crew area. The ascent
cabin has an access port at its bottom that connects directly to the top
of the surface habitat. This approach allows the cylindrical habitat to
take all launch and landing loads in the axial direction. For cargo-
only missions, the ascent stage is omitted and a much heavier habitat
can be delivered.
Another drop-tank configuration is shown in Fig. 17. LOI tanks
are dropped in LLO. After orbital capture, the CEV backs away from
the lander, the empty LOI tanks are jettisoned, and the CEV redocks.
The crew transfers to the lander and the CEV undocks and remains in
LLO. Another variant of this concept jettisons tanks that also hold a
portion of the descent propellant partway through the descent (in a
lower phasing orbit). In this case, there is a tunnel between the CEV
and the lander for crew transfer in LLO, or EVA is used.
Aside-loaded configuration is shown in Fig. 18. Surface payloads
are divided into two equal mass modules and are carried on either
side of a descent stage with central-bottom-placed main engines.
After landing, these modules can be downloaded without the use of
Fig. 11 Top-loaded lander, storable-propellant ascent.
Fig. 12 Top-loaded cargo-mission lander with surface habitat.
Fig. 13 Undercarriage concept 1, center engine.
Fig. 14 Undercarriage concept 2 with side engines.
Fig. 15 Undercarriage concept 2 with LOI tankage
Fig. 16 Drop-tank configuration 1: LOI tanks dropped in orbit: crew and cargo mission.
8DONAHUE ET AL.

dedicated cranes. In some variants, the side modules may be tall and
have a hinge at the bottom; once landed, the modules are rotated
about the hinge point into a horizontal position for placement onto a
wheeled transporter. Modules are then taken to the base for use.
Fixed-habitat configurations are not intended to have the surface
habitat removed. Figures 19 and 20 show vehicles that have descent
tankage that holds all LOI and descent propellant together. In Fig. 19,
a habitat is shown at the bottom on one side; on the opposite side is a
corresponding module, or a container holding consumables, spares,
and science equipment. Landers such as this may be intended for
multiple reuse; after an initial mission, returning crews landing
nearby return to this habitat or rely on it as a backup in case their
primary habitat fails. After several landings in close proximity,
multiple habitats would be available (though not physically linked),
and through the use of these “logistically coupled”modules, long-
duration surface missions might be undertaken.
Colony base buildup proceeds from this initial outpost
formulation, in which inflatable structures might be added,
eventually leading to a base made up of both fixed, inflatable, and
moveable assets. In Fig. 20, a spherical habitat is shown at the bottom
of the vehicle. Single H2 and O2 tanks are at the top. The descent
engines are at the bottom sides of the habitat. Engine exhausts shields
are shown. The spherical habitat pressure vessel may be nearly
identical to one of the propellant tanks and come off the same
assembly line.
Ahinged-vertical-payload configuration, shown in Fig. 21, has
descent tankage that surrounds the centrally located payload on all
but one side. In the center, above the descent engine, is a void (central
core) in which a cylindrical payload is positioned. Once landed, the
payload is rotated down to a horizontal position. A surface
transporter drives under it and transports it away to be joined to other
modules, if necessary. There is a hinge mechanism, or hinge cradle,
that attaches at the payload bottom that is designed for a one-sixth-
gravity operation. This configuration allows the payload cylinder to
take all launch, transfer, and landing acceleration loads in a vertical
direction along its axis. An ascent stage is located in the central core
for crew missions, and thus both cargo and crew versions would use
identical descent stages.
Ahorizontally integrated configuration, shown in Fig. 22, has
mirror-image descent tankage on the ends, with an ascent stage in the
center. The vehicle has a cylindrical ascent cab in the center (with
ascent engines underneath). In this case, the descent tankage,
engines, and surface air lock are left on the surface. One variation
provides for the jettison of the descent tanks on the ends before
Fig. 17 Drop-tank configuration 2: LOI tanks jettisoned after lunar
capture: crew mission.
Fig. 18 Side-loaded cargo-mission configuration.
Fig. 20 Fixed-habitat concept 2 with LOI/descent tanks at the top,
engines at the side: cargo mission.
Fig. 21 Hinged vertical payload lander: cargo mission.
Fig. 19 Fixed-habitat configuration with LOI-descent tanks at the
sides, center engine version: crew cargo mission.
Fig. 22 Horizontally integrated lander: crew mission.
DONAHUE ET AL. 9

touchdown to reduce the landed mass. In that case, the terminal
descent and touchdown propellant is drawn from the ascent tanks.
Adescent-staged configuration is shown in Fig. 23. Its combined
LOI and descent stage (CLDS) is very similar to a Delta-IV or
Centaur upper stage, with the addition of an extra engine to provide
an engine-out capability. The CLDS allows the terminal descent–
ascent stage to be much-reduced in size compared with the reference
lander. Deep throttling is no longer required by either stage, and the
modest propellant required for terminal descent is held in the ascent
tanks. The terminal descent–ascent stage uses a pressure-fed
storable-propellant engine of 6-klbf thrust. The crew has a short path
to the surface.
A terminal descent–ascent stage is shown in Fig. 24. It is
characterized by a side-positioned ascent crew cabin, central
tankage, and bottom-center engine placement. This vehicle separates
from the CLDS and descends to touchdown. After the surface
mission, it ascends to LLO.
A side-loaded concept is shown in Fig. 25. It is characterized by
side-cargo modules. These modules are hinged and are set to the top
position to fit within the launch vehicle payload shroud. While in
LLO, the modules are rotated down to a fixed and locked side
position before the vehicle descends. A crew cabin is shown on top.
This illustration is taken from [13].
An all-cryogenic O2=H2lander is shown in Fig. 26. This
illustration shows a descent stage that is common for the piloted and
cargo-only variants. Shown on the left of Fig. 26 is the piloted
version with a cryogenic O2=H2ascent stage. This stage sits within a
central void in the descent stage. The engines, located at the bottom
of the ascent stage, are used for the descent, ascent, and TEI burns.
During descent, propellant is routed to the engines from the descent-
stage tanks. On the right of Fig. 26, the cargo-only version is shown.
This illustration is taken from [13].
Fig. 23 Descent-staging lander: crew mission.
Fig. 24 Side-cabin terminal descent–ascent stage
Fig. 25 Side-cargo type-2 concept.
Fig. 26 Cryogenic O2=H2ascent lander.
10 DONAHUE ET AL.

A single-stage reusable concept that is refueled on the surface with
propellant provided by an in situ propellant plant is shown in Fig. 27.
Operating between LLO and the surface, this vehicle receives a
payload in LLO and descends to the surface. After payload
offloading, the vehicle is refueled for its next mission. This design
features a centrally located ascent crew cabin, flanked on both sides
by propellant tankage. Engines are positioned for ease of access and
change out on the surface. The tanks are located in positions that
allow for easy inspection; refueling interconnects are easily
accessed. In Fig. 27, several structural elements are excluded from
the illustration for clarity.
Conclusions
Seen from a logistical viewpoint, those lander types that allow
cargo offloading without the use of cranes may be preferred for base-
buildup scenarios. For scenarios in which the joining of habitats is
not desired, the close grouping of fixed-habitat landers may be
preferred. Most of the configurations discussed in this section do not
preclude descent-staging or drop-tank-staging modes. Boeing is
continuing its work in this area, focusing its lander concept
refinement in concert with lunar surface architecture studies. For
additional Boeing published material, see the five-year Space
Transfer Concepts and Analysis for Exploration Missions
(STCAEM) 1989–1994 reports, such as [1–4]; for Boeing material
specifically on lander design, see [5–12]. This preliminary design
study is part of a larger effort to define the attributes of future lander
vehicles and is only partially complete. The following summary
statements are preliminary and will be revised as the work progresses
and as NASA refines both its architecture and surface-asset plans.
1) The ESAS-Ares-V architecture will place, via the reference
ESAS lander, in addition to a 2.7-t ascent cabin, a surface payload of
about 4.7 t, given a 45-t lander, and a 20-t CEV is injected into lunar
trajectory by the Ares-V EDS. If the lander is 40 t, the surface payload
is about 2.8 t.
2) If maximizing landed payload is paramount, then the descent-
staging approach provides a significant advantage over the ESAS
report approach. Its other benefits (reducing the size of the lander and
reducing the final-approach deep-throttling requirement) are also
significant, though descent staging requires an engine start in the
final portion of the descent and creates a “graveyard”of used stages.
Other factors must be accounted for: for example, a complete risk
analysis, which was not done in this study, would be necessary.
Drop-tank staging also provides a modest payload advantage
without the disadvantage of a time-critical engine start.
3) The ascent-propulsion analysis indicates that the oxygen/
methane choice provides a slight surface-payload-delivery
advantage over the storable choice. Though O2=CH4Isp is 31 s
higher than storable propellant, its fuel is less dense and more thermal
conditioning would be required. More work needs to be done in this
area, and recent O2=CH4engines tests, sponsored by NASA Johnson
Spaceflight Center, are encouraging.
4) Lander configuration choice is linked to the lander
requirements, including its intended participation in the surface
architecture. If a base is to be constructed by the joining of habitats,
the undercarriage or side-loaded concepts may be preferable,
because they offer an easier pathway to site buildup. Habitats
preintegrated to wheeled transports could be moved to the base
immediately after landing without the additional steps of placing a
dedicated crane on the surface and conducting a host-off maneuver.
Side-loaded configurations require the main cargo to be divided but
to retain bottom-center engine placement. If offloading and joining of
habitats is not an objective or if the offloading and joining is to be
performed with the aid of preemplaced surface equipment, then the
top-loaded concept may be preferred.
Acknowledgments
The authors would like to thank Mike Lounge, Keith Reiley,
Curtis Maulsby, and Ben Barackman.
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Associate Editor
Fig. 27 In situ propellant single-stage lander.
DONAHUE ET AL. 11