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Global Space Exploration Conference (GLEX 2017), Beijing, China 6-8 June 2017.

GLEX–2017–3.2A.1

Global Exploration Roadmap Derived Concept forHuman Exploration of the Moon

Ryan Whitley a, Markus Landgraf b, Naoki Sato c, Martin Picard d, Kandyce Goodliff e, KeithStephenson f , Shinichiro Narita g, Yves Gonthier h, Aidan Cowley i, Shahrzad Hosseini j k,

Rogier Schonenborg l

aNational Aeronautics and Space Administration, United States, [email protected] Space Agency, ESTEC, The Netherlands, [email protected] Aerospace Exploration Agency, Japan, [email protected] Space Agency, [email protected] Aeronautics and Space Administration, United States, [email protected] Space Agency, ESTEC, The Netherlands, [email protected] Aerospace Exploration Agency, Japan, [email protected] Space Agency, [email protected] Space Agency, EAC, Germany, [email protected] Space Agency, ESTEC, The Netherlands, [email protected] de l’Aeronautique et de l’Espace (ISAE)/Supaero, FrancelEuropean Space Agency, ESTEC, The Netherlands, [email protected]

AbstractTaking advantage of the development of Mars-forward assets in cislunar space, a human lunar surface conceptis proposed to maximize value for both lunar exploration and future deep space missions. The human lunarsurface missions will be designed to build upon the cislunar activities that precede them, providing experiencein planetary surface operations that cannot be obtained in cislunar space. To enable a five-mission limitedcampaign to the surface of the Moon, two new elements are required: a human lunar lander and a mobilesurface habitat. The human lunar lander will have been developed throughout the cislunar phase from a sub-scale demonstrator and will consist of a descent module alongside a reusable ascent module. The reusableascent module will be used for all five human lunar surface missions. Surface habitation, in the form of twosmall pressurized rovers, will enable 4 crew to spend up to 42 days on the lunar surface.

1. Introduction

The Global Exploration Roadmap (GER)1 is a doc-ument published by a group of international spaceagencies, formally known as the International SpaceExploration Coordination Group (ISECG), whichidentifies opportunities to unite national plans into amulti-faceted collaborative framework. The roadmapendeavors to maximize the contributions of each inter-national agency to simultaneously meet well-definedglobal exploration goals, a truly win-win propositionfor the participants. With exploration destinationsranging from low Earth orbit and International SpaceStation (ISS) to the Moon and Mars, the GER hasthe potential to engender collaboration in all aspectsof future human space exploration.

In this paper, opportunities for collaboration thatlead to human return to the lunar surface are postu-lated. Most recently, NASA has proposed the devel-opment of a Deep Space Gateway (DSG) around theMoon that will provide the ability to support multipleU.S. and international partner objectives.2 The DSGconcept, described as the Evolvable Deep Space Habi-tat in the GER, is emerging as the consensus next step

that meets multiple exploration paths including theMoon, Mars and beyond. The DSG will be built upover a series of launches, both as crew co-manifestedpayloads and independent smaller launches, and willcontain functions that permit habitation, EVAs, as-sembly, science experiments, and technology demon-strations while serving as a docking station for visitingvehicles. Contributions from all nations are currentlybeing welcomed that will directly lead to future deepspace exploration missions.

To that end, a cislunar orbit for the DSG has beenselected that represents a bold step beyond Low EarthOrbit, is reasonably accessible from Earth-based space-craft currently available or in development, all whilemaximizing the ability for all nations to conduct mis-sions from a common location to multiple deep spacedestinations. Known as the Near Rectilinear Halo Or-bit or NRHO, it is a 3-body halo orbit around theEarth-Moon Libration Point 2 that exists sufficientlyout of the lunar gravity well to be readily accessiblefrom Earth-based spacecraft, but with properties fa-vorable for relatively fast and cheap transits to thelunar surface compared to other comparable cislunarorbits.3

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Table 1: Mapping Lunar Surface Principles

GER Principle Surface Campaign Principle

AffordabilityLimit lunar surface infrastructure to what is necessary to achieve priority lunar

exploration goals

Exploration ValueUse unique value of astronauts on the lunar surface and in lunar orbit foradvancing goals and objectives of lunar, fundamental, and applied science

International PartnershipsDefine a modular surface architecture, which maximizes opportunities for

international partnership

Capability EvolutionPrepare for Mars surface exploration and maximize synergies between lunar and

Mars surface campaigns

Human-RoboticPartnership

Identify opportunities for robotic elements to augment achievement of goals andensure safety of the astronauts

RobustnessProvide dissimilar redundancy for critical functions on the lunar surface and

define contingency/abort modes

The return journey to the lunar surface will beginfrom the Deep Space Gateway. How would the DSG beused to support human lunar surface missions? Someoptions include: storage and refueling of a reusableascent module, safe haven for crew aborting from thesurface, communications relay to Earth from the lunarfar side, teleoperations of robotic systems (on the farside) as well as access to medical and exercise equip-ment for reconditioning when returning from a lunarsurface mission. Probably the largest benefit to stagingfrom the DSG is the potential to enable partial landerreusability. The DSG could become a hub for refuelingboth crew and spacecraft consumables which in turncould help open new commercial markets. The trans-portation trade space is examined fully in the nextsection where it is demonstrated that the DSG staginglocation fits favorably within a lunar surface architec-ture.

2. Mission Architecture

Knowing that the number of possible lunar surfacearchitecture perturbations are endless, the consensusGER lunar surface mission concept was narrowed byconsidering three primary inputs: 1. GER derivedstrategic principles, 2. GER derived goals and objec-tives and 3. Capability based constraints as framed bypotential international contributions.

2.1 Architecture Drivers

In the case of the first guideline, the ISECG com-munity crafted a set of lunar strategic principles bymapping them directly from previously published GERprinciples1 as shown in Table 1.

Of these principles the two that were most critical indetermining the mission architecture were the desire topromote “Affordability” and “International Partner-ships”. The combined principles could be summarizedas an endeavor to maximize partnership opportunitiesby prioritizing modular systems while at the same time

minimizing cost and complexity, ultimately favoringminimum mass solutions.

The second input to the study was a list of clearlystated goals and objectives, articulating that the hu-man lunar return should be both focused and fresh.The primary architecture drivers were specified sim-ply: the lunar surface campaign should feature 5 mis-sions of 28+ days each and enable crews of 4 to reachthe lunar surface at any one time. A more detailedbreakdown of individual objectives, based on the GERgoals and objectives were identified to drive technologydevelopment programs and address priority scientificopportunities.

Finally, the various architectures must align withthe capabilities currently available or in developmentby the international partners engaged in the humanlunar return activity. At this point, the architec-ture features involvement from the Canadian SpaceAgency (CSA), the European Space Agency (ESA),the Japanese Aerospace Exploration Agency (JAXA)and the United States National Aeronautics and SpaceAdministration (NASA). As such, the envisaged ar-chitecture features expertise and capabilities availablefrom these communities, including the active launchvehicles from ESA (Ariane V) and JAXA (H-II) aswell as the crew transportation system currently beingdeveloped by NASA (Space Launch System (SLS) andOrion).

2.2 Trade Space Options

There are two primary aspects of a lunar surfacearchitecture: the lunar lander design and the numberand type of surface elements. The ISECG commu-nity accomplished the first aspect of the architectureby building a trade matrix featuring potential staginglocations, launch vehicles, and the number and typeof propulsive elements. All architectures must neces-sarily adhere to the dynamics governing the mechanicsof spaceflight. Thus, trading the staging locations is

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Fig. 1: Staging orbit options and costs

especially important for determining the total propel-lant cost (∆V ) required, identifying which element willperform each burn and understanding the active life-time of the crewed module. Each aspect is critical indetermining design feasibility.

To that end, Figure 1 provides a breakdown ofpropulsive costs and transfer times associated with arepresentative set of staging orbits: Distant Retro-grade Orbit (DRO), Earth-Moon L2 Halo, Near Rec-tilinear Halo Orbit (NRHO) and Low Lunar Orbit(LLO). While other orbits exist in between the fourshown in Figure 1, the examples cover a large swathof orbits both close to the moon such as LLO or theDRO which is far from the moon. Thus, the represen-tative orbits are useful proxies for determining propul-sive requirements and crew systems / consumables tobe levied on a lander design.

The preferred orbit from a lander standpoint wouldbe LLO. However, previous studies3 have demon-strated that LLO (as well as other low lunar energyorbits similar to LLO) require propellant loads thatexceed NASA’s Orion vehicle capabilities. One optionwould be to add an in-space propulsion stage to theSLS/Orion architecture to enable LLO and then de-sign a lander to operate out of LLO. For this study theassumption was made to limit the number of new ele-ments to the lander and surface based systems whichconstrains the staging orbit to an orbit larger thanLLO.

Comparing the three remaining orbits reveals a bigdifference between NRHO and the other orbits, notin terms of total ∆V , but in terms of vehicle lifetimeand crew consumables required. A lunar lander has

what is called a very large gear ratio in terms of massof propellant required per unit of dry mass. This isdue to the fact that a lander must only use thrustersto descend to the lunar surface without atmosphericdrag to use as a decelerator. The total cost is about2,000 m/s each way or 4,000 m/s total. Depending onthe design a two stage lander could require up to 10kg of propellant for every 1 kg of additional payload.Thus, it is critical to minimize the systems dry mass,especially systems necessary for the crew such as con-sumables. As Figure 1 demonstrates, the DRO and L2Halo orbits have round trip time of 8 days and 6 daysrespectively. This is true for most higher energy cis-lunar orbits. The NRHO on the other hand is uniquedue its once per revolution close perilune passage andthus can do the round trip within 1 day (1/2 day eachway) one time every orbit period. That is equivalentto a 85% to 90% reduction in crew time on the landerwhich is directly correlated to decrease in the mass ofsupplies, air and associated systems that result in amultiplied increase in propellant mass. Thus, with the∆V costs relatively comparable, the NRHO is the pre-ferred staging orbit of the 3 cislunar orbits examined.

The remaining trade matrix items for the lander de-sign include propellant type and staging strategy alongwith launch vehicle specifications. The biggest massdriver for the lander trade study is the propellant typeof the descent module. While the propellant efficiencyof the ascent module is important, it was much less of afactor than the propellant type of the descent modulewhich must transport the ascent module as payload.As a result, a high Technical Readiness Level (TRL)storable bi-propellant combination (MON/MMH) was

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Fig. 2: Potential combinations of propellant type and staging strategy

selected for the ascent module. The transportationtrade tree then reduces to two items, 1) the propellanttype of the descent module and 2) the staging strat-egy. If the launch vehicle is considered as a constraintrather than a trade variable, the remaining trade de-cisions can be broken down as shown in Figure 2.

To that end, the feasibility of various human lu-nar lander propellant types in the transfer scenariowere assessed considering three primary thruster types:LO2/LH2, LO2/CH4, and MON/MMH. Performanceassumptions of each main engine are given in Table2 below. The primary concern of the LO2/LH2 andLO2/CH4 propellant options is boil-off. Liquid hydro-gen is also problematic from a volumetric standpoint.While technology projects are in work to mitigate thepropellant loss due to boil-off, technology solutions arelow TRL. Thus, in this analysis conservative figures as-sociated with boil-off were applied.

Table 2: Assumptions for Propellant Trade Study

Item Value

∆V to Surface 2,800 m/sLEO to DSG Weak Stability Boundary Trans.Transfer Time 100 daysPayload Mass 10,000 kg

LO2/LH2 LO2/CH4 MON/MMHBoil-Off Rate 1.0% 0.1% 0.0%

Isp 450 s 370 s 320 sStruct. Ratio 25% 20% 15%

Trade-off results are given in Table 3 assuming a 10t payload to the surface (which is a few tons lighterthan the current ascent module design, see Section 3).Due to the high boil-off mass the total human lunarlander mass with an LO2/LH2 descent module is over55 t at the launch time. This mass significantly ex-

Table 3: Descent Module Propellant Trade-off Results

Item Value

LO2/LH2 LO2/CH4 MON/MMHFuel Mass 34.0 t 18.1 t 19.3 t

(boil-off mass) 21.4 t 1.7 t 0.0 tDry Mass 21.3 t 14.5 t 13.4 t

Total Mass 55.3 t 32.6 t 32.7 tLunar ISRU All Prop O2 Only NoneMars ISRU All Prop All Prop None

ceeds the Trans-Lunar Injection (TLI) capability of theSLS Block 1B Cargo vehicle. A main engine that usesLO2/CH4 fuel has a significantly lower boil-off ratecompared to a LO2/LH2 engine, a decrease of about90%. Thus, the total lander mass for a LO2/CH4 stagedescent module decreases to about 32.6 t. It is theminimum mass result of these three options, and isfeasible for a SLS Block 1B Cargo launch, even whenscaled up to the actual ascent module size closer to 14t. In addition, this engine can take advantage of In-Situ Resource Utilization (ISRU) scenarios for futureexploration at both the moon and Mars. A main en-gine featuring storable MON/MMH fuels have no boil-off gas and has the highest TRL. The mass of storabledescent module is very close to the LO2/CH4 enginecase as the Isp difference, boil-off mass and dry massdifferences tend to cancel each other out. However, astorable engine does not lend itself to ISRU options,and there is an opportunity for further LO2/CH4 massdecreases if the TRL of boil-off mitigations advancefurther. Thus, LO2/CH4 was selected as the referencedescent module propellant. In addition, the more tra-ditional staging strategy was selected as the reference,recognizing that a drop stage has advantages of in-creasing some level of reusability and minimizing hard-

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Fig. 3: Final Trade Space

ware on the lunar surface. However a drop stage makescargo delivery more difficult as a descent module wouldrequire significant modification or a new element alto-gether.

With the reference transportation elements set, amatrix of surface system options were identified. Thefinal version of the trade space, both options and refer-ence selections, is shown in Figure 3. The primary de-sign choices were the Extra Vehicular Activity (EVA)method, surface habitation strategy, ISRU approach,and power system. The level of exploration activitiesto be conducted during the lunar night played a largerole in design choices and overall mass. Three poten-tial levels of exploration activity were identified: nolunar night operations, limited lunar night operations,and full exploration lunar night operations. These lev-els were then juxtaposed with the GER principles andguidelines to seek a diverse set of landing sites in asingle region so that all 5 missions were successful inachieving unique scientific value. As a result, a mo-bile habitat with greater than 30 day stay capabilitywas selected as the preferred implementation for thecrewed stays. The other option, a stationary habitat,would require deploying a new habitat at each poten-tial new exploration zone. In contrast, a mobile andreusable habitat would require only one new element(or two copies of the same element but delivered onthe same cargo lander).

The most interesting trade relates to the strategy forlunar night operations. If a 14 Earth day lunar nightcycle is to be achieved, the power system must notrely extensively on solar-panels and batteries. Surviv-ing the lunar night using only batteries requires bat-

teries that are so big that they exceed the allocatedmass/size budget. The only other way to conduct a28+ day mission without 14 days of darkness is to parkon a permanently lit region at one of the poles. How-ever, this conflicts with the goals and objectives of themission to explore diverse regions. So the trade is re-duced to two options: full night time vs limited nighttime operations. In the end, the design focused on lim-ited operations as a full night time the power system,either nuclear or fuel cell or a combination of both wastoo heavy to meet the mass constraints of the lunarlander cargo design. A nuclear-based (radio-isotope)system supplemented by solar panels and batteries wasselected for the reference design (described in more de-tail in Section 3). In addition ISRU requirements wereeliminated due to the risk of both quantifying avail-able resources (part of the reason for returning to themoon) and the methods of extraction required.

3. Lunar Exploration Elements

As discussed in the Section 2, the concept of thehuman lunar lander is driven by the goals and objec-tives of lunar exploration in the frame of the globalexploration roadmap and informed by the associatedstrategic principles given in this paper. Guided by thatframework, the human lunar lander has been concep-tualized to be a two-stage, partially reusable vehiclefor four crew providing mainly transportation betweenDSG and the surface, and only short-term habitation.The surface habitation function is to be provided bythe pressurized rover. The lander mission is designedto protect for global access to the surface, however, not

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Fig. 4: Envisioned co-development approach. The time scale on the top of the figure indicates years to HumanLunar Return.

allowing this requirement to impact the affordability ofthe design. Further, the descent stage of the humanlander must double as a large cargo lander.

The human lunar lander is sized for a single SLSBlock 1B Cargo launch but remains compatible witha dual launch or as a co-manifested payload (CMP)partially fueled. The lander arrives and rendezvouseswith the Deep Space Gateway (DSG) and loiters untilthe crew arrives on a separate launch. After docking,the crew transfers via the DSG to the lander. Thelander carries crew to the lunar surface and returnscrew and samples to the DSG after surface operations.

The reference configuration features a two-stage lan-der comprised of a descent module (DM) and ascentmodule (AM). The DM completes all descent maneu-vers from the DSG until touchdown. The DM is lefton the surface of the moon while the AM returns tothe DSG for reuse. The cargo variant of the landerdoes not carry an AM. There is also an optional con-figuration with a drop stage. The design figure of thetwo-stage lander is shown in Figure 6.

3.1 Element Co-Development Plan

A programmatic analysis was performed of howsuch an architecture could be realized. The study find-ings indicate that the approach to develop a sub-scalerobotic demonstrator lander allowing the implementa-tion of system components (see Figure 4) correspond-ing to the partnership roles in the human architectureis the most affordable solution for the international co-ordination.

The timing of the programmatic realization is con-strained by the assumed decision making process, theduration of the development of key technologies, thetypical time scales of space vehicle design and inte-gration, and the relative timing of the demonstrator

mission implementation with respect to the human ar-chitecture. The programmatic analysis has found that:(a) the time between the launch of the demonstratormission should be between 4 and 6 years before hu-man lunar return, and that (b) the introduction of ademonstrator mission does not significantly increasethe duration of the end-to-end time from decision tohuman lunar return.

Regarding finding (a) it should be noted that timesshorter than 4 years would not allow sufficient impactof findings of the demonstrator mission to influencethe design and production of the human mission, andthat times longer than 6 years risk the demonstratortechnologies becoming obsolete for the human archi-tecture. Finding (b) is based on the assumption thatkey technologies and components (e.g. descent propul-sion) drive the development of the human architectureand that identical components are used on the demon-strator and the human vehicles. Without a demonstra-tor, the human architecture would still require flight-demonstration in order to achieve human rating. Itshould be noted that with different programmatic as-sumptions, the results can be different.

The conceptual designs for the elements of therobotic demonstrator and the human lunar lander andthe demonstrator mission are described in more detailbelow (see Figure 5).

3.2 Reusable Human-Rated Lunar Ascent Module

Industrial concept studies performed earlier thisdecade have yielded the basic configuration and sys-tem characteristics of a reusable ascent module hostinga crew of four for three to four days of transportation.The main function of the ascent module is to deliverthe crew and samples safely to the DSG after the endof the surface mission. Its habitation function is re-

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Fig. 5: Relationship between demonstrator and human elements

quired also for the descent and in a short contingencycase, in which the crew cannot reach the pressurizedrovers. Further, the ascent module has to support thevarious contingency and abort modes that allow safereturn of the crew to the DSG at any point along themission (this includes, for example, the return fromthe lunar surface during night).

On-going concept studies of the ascent module havethe objective to further reduce the overall mass per-formance of the ascent module taking into account thesignificant mass leverage of the system dry mass withrespect to the vehicle overall wet mass (as applicable toits mission initial state at the DSG) of 1:7.3. The basicidea is to tailor the concept for the mobility-based mis-sion scenario with relatively short time the crew spendsinside the ascent module (assumed two times 12 hoursfor transfer plus EVA preparation). The concept reliesthus on a minimization of the structural mass of thevehicle through reduction of the pressurized volumethat is available to the crew from a previously set 20m3 to 10 m3. Besides improving the mass performanceof the concept, the volume reduction also enables con-sideration of alternative split-launch concepts with awider variety of launch vehicles.

In order to achieve a good mass performance, theatmospheric pressure inside the ascent module will bebelow standard atmosphere. This is in principle com-patible with the standard atmosphere inside the DSGassuming the presence of an airlock with docking portsfor the human lunar lander. Besides the reduction instructural mass the lower pressure inside the ascentmodule allows reducing the mass of the life supportsystem.

For propulsion of the ascent module two separatesystems have been conceptualized. The main ∆V willbe provided by either a single large or a cluster ofsmaller, pump fed engines operating at a mixture ra-tio chosen for maximum performance in specific im-pulse. A pump fed engine has a higher chamber pres-sure, which allows for a more compact and lower massengine, while at the same time achieving a higher spe-

Fig. 6: Human Lunar Lander Schematic

cific impulse than a pressure fed engine. At the sametime lower subsystem masses can be achieved throughlower tank pressures. The engine cluster option comeswith the added advantage of offering a flight demon-stration of an identical single engine in the frame ofthe sub-scale demonstrator.

Other aspects of the ascent module that is cur-rently under study are the opportunities for a com-mon human-machine interface shared between the as-cent module (crew control of landing and ascent) andthe robotic control terminal in the DSG (monitoring ofvisiting vehicles and robotic capture and berthing, andtele-operations of surface assets such as the demonstra-tor rover).

3.3 Human-Rated LOX/CH4 Descent Module

As described in Section 2, LO2/CH4 was selectedas the propulsion system for the descent module. Forthe majority of the mission, the ascent module is theprimary source of avionics and vehicle control, whilethe descent module has avionics necessary for inde-pendent operation as well as in the cargo variant. The

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AM is reusable, while the DM is not. Both AM andDM elements provide power to the human lander asa whole. In addition, while waiting for the crew, theDSG provides power for the system while the landerremains attached. After separation, lander uses powerfrom AM and DM. Key features of descent module aredescribed in Table 4.

Table 4: Key Features of Descent Module

Item Descent Module (DM)Main Engine Propellants LO2/CH4

# Main Engine Casters 3Main Engine Thrust >80 kN

Main Engine Isp 370 sRCS Thrusters 40 x 220 N

Power Body Mounted Solar Panels

3.4 Reusable Surface Rovers

Following the surface access trades conducted, theinternational team iterated on a conceptual design ofthe Small Pressurized Rover (SPR) for crew habitationand mobile exploration of the lunar surface. Amongstthe different constraints, a reduced number were driv-ing significantly the pressurized rover concept. Givena surface mission of 42 days for a team of 2 up to4 crew members per rover in a contingency situation,the habitable volume had to be sufficient to fulfill crewneeds. This added to the constraint of launching thetwo rovers on the same SLS flight and having the roverconfigured with an airlock and most likely radioiso-topes sources for power and thermal resulted in somevery interesting challenges. After some iterations andtrade studies, the concept evolved and matured intoa notional representation based on a vertical launchconfiguration of the two rovers as illustrated in Fig-ure 7. This configuration option would be to launchboth rovers back to back resting on a vertical attach-ment structure that could also serve as the deploymentsystem as illustrated in Figure 8.

Fig. 7: Notional launch configuration of rovers

The rover itself can be broken down into four mainsystems, opening the door to partnership and sharingof the pressurized rover development:

a Mobility and cross-cutting sub-systems, essen-tially the rover platform

b The radioisotope power/thermal source

c The airlock, and

d The pressurized module

At this point in the rover design process, mass es-timates for the airlock and the pressurized portions ofthe element can be calculated by simply scaling ex-isting hardware elements and/or other mature designplans. The module that needs further design consider-ation due to its uniqueness is the design of the secondmain system listed: the radioisotope power/thermalsource.

The envisaged hybrid power system includes par-tially deployable solar panels, multiple radioisotopepower systems and an electrochemical energy storagesystem of either rechargeable batteries or regenerativefuel cells. The design and sizing of such a system is acomplex multivariable problem, with the optimum ar-chitecture being dependent upon the operational con-cept, the power requirements of the various electricalsystems required to deliver those operations, the lunarsurface environmental and lighting conditions, and theperformance characteristics of the various power sys-tem technologies.

A complex parametric model of a lunar rover powersystem was created in order to investigate the varioustrade-offs, determine the optimum configuration andestimate the power system mass. A meaningful powersystem architecture for the rover could be determinedonly by considering, in parallel, some key aspects ofthe thermal management system. The cooling require-ments of the rover during the lunar day demand thatmuch of the available body area is dedicated to radia-tors, and is therefore unavailable for solar power gen-eration. During the night, heating power may form alarge proportion of the electrical load requirement.

An adequate autonomous power system was found,dependent on the availability of radioisotope powerunits that enable the lunar-night survival. The util-isation of compact radioisotope systems also allevi-ates the difficulty balancing the daylight energy bud-get given the constrained solar panel area. A well-insulated rover with a switchable fluid loop cooling sys-tem should keep the crew warm in the lunar night withlittle heating power above that provided by the elec-trical systems and human metabolism. Any additionalthermal support would be most effectively provided bylarge radioisotope heater units.

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Fig. 8: Notional deployment concept for rovers

4. Surface Operations Concept

Armed with a reusable pressurized rover design, arobust surface operations concept is devised. Withan exploration blueprint consisting of five consecutivemissions and a plan to reuse the pressurized roversfrom mission to mission, an operations concept mustboth examine how to conduct a 42 day crewed sur-face stay with lengthy light and night cycles as well asfind scientifically interesting landing sites that are closeenough to each other to enable telerobotic relocationof the rovers between missions.

Fig. 9: Notional Landing Sites

4.1 Landing Sites to Connect 5 Missions

There are multiple regions on the Moon where sucha set of five missions enables significant regional ex-ploration campaign to be undertaken. Examples in-clude both the North and South Poles, multiple im-pact craters, and volcanic features such as lava flows.Based on the consolidated, globally recognized ratio-nale for human exploration of the Moon4,5 a sequenceleading from the South Pole Region to the interior ofthe South Polar Aitken Basin (SPAB) has been identi-

fied as a good reference. The South Pole Aitken basinis the largest impact structure on the Moon, indeedone of the largest in the Solar System. With a diame-ter of approximately 2500 km, it dominates the lunarfarside, extending from the South Pole at one pointon its rim to the 135 km Aitken crater at the oppo-site point on the basin rim. This exploration region iscompatible with an orbiting staging location (visibilityfor communication relay and transfer ∆V ) of the DSGon an NRHO, and it allows addressing the top-priorityscience and exploration goals.

Table 5: Coordinates of notional landing sites

# Site Lon. Lat.1 Malapert Massif 0 -852 Shackleton Plateau 126 -893 Schrodinger Basin 139 -754 Antoniadi 172 -705 SPAB Interior 160 -60

Even just considering the South Pole Aitken basinas a place for humans to explore, there are many morethan five interesting potential landing sites. For thisstudy five notional sites have been chosen that extendfrom the rim of the basin to the far interior of the basinas shown in Figure 9. The first two are both on therim of the basin, Mallapert Massif, and the South Pole(located on the rim of the 20 km Shackleton crater).Then moving into SPAB there is Schrodinger basin,the Antoniadi crater (which contains the lowest eleva-tion on the Moon), and finally arriving at a point closeto the center of the SPAB. Coordinates for all five aregiven in Table 5. The five sites represent a good designcase for the architecture concept driving power, ther-mal, and communication subsystem designs, as well asthe assumptions on the terrain environment for surfacelanding and mobility. For each potential site, thereare many opportunities for science and future explo-ration. Some of the major science and future explo-ration themes at SPAB are described in Table 6. Adetailed breakdown of notional SPAB landing sites is

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Table 6: Science and exploration themes at the SPAB

Theme ObjectiveGeological Features Geological Exploration as wide area of SPAB

Water or Ice Lunar Water/Ice or Volatile component exploration for future in-situ resource utilization.Observation Moonquake observation. Astronomical Observatory on lunar surface.

Fig. 10: Example landing sites within the SPAB

given in Figure 10.

4.2 Crew Concept of Operations

The surface operations of the human architecturerely entirely on the capabilities of the two pressur-ized rovers, each of which provide habitation for twocrewmembers in the nominal case. In order to ad-dress the envisioned science goals and to gather therequired operational experience for future planetarysurface exploration (e.g. Mars), the surface missionduration cannot be limited in general to a single lu-nar day. While the earlier missions in the programmight be shorter, in general the mission duration isenvisioned to be constrained by a day-night-day cycleequal to 42 days, requiring night survival for the crewand equipment. The operational modes for lunar dayand lunar night will be significantly different. Whileduring daytime the crew drives to multiple scientifi-cally interesting sites for investigation and sampling,they return to the vicinity of the lander for the nightand spend the night inside the rover with analysis ofprevious EVAs and incremental training preparing forthe next lunar day. This approach will not only enablea feasible power and thermal design for the pressur-ized rover, but also enable a safe abort scenario duringlunar night. Another key operational concept is to ex-ploit the split of the crew between rovers “Crew A” and“Crew B”) with an approach of alternating schedule ofrover-based EVA every second day for each crew. This

would facilitate the coordination with mission controland science backroom on the ground.

Contingency operations are driven by the goal ofkeeping the crew safe. At any point in time the crewwill be able to return to the DSG within 36 (TBC)hours, out of which 24 (TBC) hours are allocated to thereturn from the crew’s location to the lander. For thisreason, the maximum distance of the rovers from thelanding site must be lower than 100 km and the roveraverage speed in the reference terrain approximately 5km/hr. During lunar night, the rovers will be parkedsufficiently close to the lander, so the crew can reachit on foot in EVA suits within 4 hours.

A more detailed chronological overview of the sur-face crew operations could proceed as follows (“CrewA’s” perspective is illustrated by Figure 11): after thelanding at lunar dawn, the crew performs an EVA fromthe lander to the rover and initiate the commissioningof the rover. Once the safe operation of the pressurizedrover is ensured, Crew A drives to the first explorationsite during the first Earth day while Crew B stays inthe vicinity of the lander. Crew A surveys the geologyat the first exploration site and prepares the rover-based EVA in cooperation with the science backroomon the ground. The two crew dons for EVA and leavethe rover through the airlock for surface exploration ofup to 4 hours. On the next Earth day Crew B repeatsthe process at a different exploration site while CrewA remains inside the rover. The crew that is active on

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Fig. 11: Concept of Operations for Human Lunar Surface Exploration: “Crew A” Perspective

the surface might complete multiple excursions duringa single shift. Each shift ends with 8 hours of crewrest. The total mobility distance during the lunar dayis several hundred kilometers.

Both crews drive back to close to the landing sitebefore the Sun sets on the first Moon day. During lu-nar night, as the solar arrays will no longer be gener-ating energy, the rovers enter a hibernation operationmode. Each Earth day during the night cycle is spentwith in-situ analysis and on-board training preparingthe next lunar day as well as performing public eventstransmitted back to earth. The second lunar day isspent similar to the first, with the exception that thereturn to the landing site occurs earlier to allow for thefinal EVA to the lander. Once inside the lander, thecrew re-pressurize the ascent stage of the lander, dofftheir suits, clean all items that were in contact withthe lunar dust environment, and start the process ofcommissioning the ascent stage of the lander for safereturn. The ascent burn is initiated after full function-ing of the ascent stage has been confirmed completingthe surface operations.

Another aspect of surface operations is the activi-ties of the rovers when crew is not present. Since thepressurized rovers have to be transferred to the nextlanding site prior to the next crew arriving, the roversmust be controlled to drive across long stretches of lu-nar terrain to reach the next landing site in sequence.A team of scientists have investigated potential tra-verses and have found less than 3000 km traverses into

the SPAB from the South Pole with maximum slopesless than 23.6 If the rover could drive at 0.1 m/s onaverage, the accumulated driving time would be a littleless than a year to cover the full distance.

4.3 Aspects of Human-Robotic Integration Operations

During the 42-day human surface mission, the crewhas various tasks to accomplish, amongst which aredriving the rover and to collect samples from the lu-nar surface. The pressurized rover is driven by theon-board crew. Sampling can be done via EVA andvia a robotic arm on the rover. During an EVA, thecrew member can approach the sample while takinginto account the safety requirements - and use manualtools to collect the samples. This can be advantageousin areas where the rover cannot easily access due to thesurface or the size of the rover, e.g. a smaller rock inbetween two larger rocks. In these scenarios, the dex-terity of the astronaut on the surface can potentiallyenable a larger variety of samples to be retrieved. Inthe second option, when the robotic arm of the rover isused, this requires a human-machine interface (HMI)aboard the rover. The crew member inside the rovercan control the robotic arm by operating it using theHMI. In case of the demonstrator mission (see separatesection below), the same operations can be performedby the crew aboard the DSG using an HMI onboardthe station. In this case, a small, yet limited, delay isintroduced, and potentially the use of force feedbackcould to enhance the control operations.

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The robotic arm on the pressurized rover is con-trolled by the crew members aboard the rover, or, al-ternatively by HMI if crew are not present, such asduring the traverse periods. Multiple end-effectors arerequired on the robotic arm in order to enable the col-lection of different types of samples. Examples of thefunctions of the end-effectors are chipping pieces ofrock(let), scooping regolith and volatiles, and takingcore samples. Another advantage of the robotic arm isits reach, especially for places that cannot be accessedby the crew, e.g. for sampling sites that are too highor high-risk for the crew to reach physically.

The surface operations of the Apollo program haveprovided valuable insight in efficient approaches tocrew surface operations. It has shown that at timesthe collection of samples or tool handling were chal-lenging for the crew at certain parts of the operations,while at other parts of the operation it was essential.The presence of both humans and robotic elements onthe surface in the reference mission scenario describedabove introduces human-robotic interaction that canbe used to enhance the efficiency of surface operations.This can for example be achieved if parts of the op-erations of the robotic element are automated for ef-ficiency, or if variable automation can be introducedsuch that the surface crew or the mission control teamcan make decisions to take over the control from therover.

4.4 Preparing for Human Lunar Return witha Sub-Scale Demonstrator

An integral part of the approach for co-developmentof the human lunar lander in an international part-nership is the preparation of technologies, operations,components, and roles in the frame of a human lunarsurface demonstration mission scenario. One findingof an internationally coordinated of Human Lunar Ex-ploration Precursor Programme (HLEPP) study,7,8 isthat a sub-scale robotic lander can achieve flight qual-ification and mission preparation in the most efficientway. At the same time, such a development will imple-ment the small cargo lander and provide opportunitiesfor significant scientific investigations of previously un-explored regions of the Moon.

Like the human lander, the robotic lander is com-prised of a descent stage, ascent module, and a rover(see Figure 5). The propulsive stages feature identi-cal rocket engines thus enabling their qualification inflight. The higher thrust requirement of the four timesheavier human lander is met by combining the enginesthat are flown as single units on the demonstrator intoclusters on the human lander. Another element that isconsidered identical or at least representative is theguidance, navigation, and control subsystem.

At the center of the operational demonstrations is

the robotic rover that will spend the first (approxi-mately 70 days) part of its mission collecting samplesfor return to the DSG, where they will be received bya crew for ultimate return to Earth. After the roboticascent module will have left the surface, the rover willcontinue to traverse the lunar surface in the 1-yearmission to demonstrate long-range durability and re-liability of planetary surface mobility. Other opera-tional demonstrations achieved by the demonstratormission are landing operations, ascent operations, andrendezvous operations with the DSG. Key technologiesrequired for enabling the human architecture are high-efficiency throttle-able propulsion for descent (e.g. liq-uid O2/CH4 bi-propellant), hybrid power systems fornight survival, and durable mechanisms for long-range,long-duration mobility in dusty environments.

Based on the objective of flight demonstration de-scribed above, a concept for the HLEPP mission sce-nario has been advanced to phase-0 level by the par-ticipating agencies CSA, ESA, and JAXA (see Fig-ure 12). A mid-sized expandable launch system withperformance in the order of ten metric tonnes into Geo-stationary Transfer Orbit (GTO) such as Ariane 5 (orfuture evolution) is sufficient to put the robotic lan-der on a transfer to the Moon (either via weak stabil-ity boundary, WSB, or minimum-energy depending onthe boil-off characteristics of the semi-cryogenic pro-pellant in the descent stage). After landing, the sur-face mission of the rover is envisioned to use a lunarday-night-day-night-day cycle to cover tens of kilome-ters to collect up to 15 kg of samples from previouslyunexplored regions of the Moon. During the surfacemission the interface to the DSG comes into play whena crew arrives. The crew tele-operates the rover thusgaining operational experience of controlling a vehicleon the lunar surface. The crew vehicle docked to theDSG plays an essential role after robotic ascent mod-ule has launched from the surface and after it has beenberthed to the DSG: the precious samples (red ovalshown in Figure 12) collected during the surface mis-sion are transferred to the crew vehicle for safe returnto Earth. The operations of the elements of the roboticdemonstrator mission thus mimic a human mission,thus improving the operational confidence and certifi-cation level of key components for the human lander.

Finally, the robotic demonstrator will be a vital partof enabling reusability of the human systems, as onlyenough supplies for a single 42 day mission will be man-ifested on the rovers to reach the mass targets. Thus, acopy of the demonstrator lander will be used to resup-ply the crew and vehicle consumables. At this point,between 1 and 2 small cargo vehicles are required toresupply with a goal to reduce to a single cargo vehicle.

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Fig. 12: Scenario of the lunar surface demonstrator mission

5. Conclusion

Work on realizing elements of the Global Explo-ration Roadmap has started to paint a picture of an af-fordable, yet valuable scenario for human lunar explo-ration based on decades of previous work. With somekey elements for human spaceflight beyond Earth or-bit coming together today - such as the Space LaunchSystem and Orion crew vehicles - as well as more andmore international partnerships and private sector ini-tiatives in lunar exploration becoming reality, now isthe time for next steps. Already today a partnershipof ISECG participating agencies is considering to moveahead in the construction of the Deep Space Gateway,which could be the staging location for, in the firstphase, a sub-scale demonstrator mission and, in thesecond phase, return of humans to the surface of theMoon. The demonstrator mission is moving beyondthe conceptual phase and into a focussed and coordi-nated design study phase with the objective to considerkey decisions by participating agencies towards the endof this decade. While the scenario described above is aproduct of a minimalist approach to conceptualization,affordability remains a major risk for the implementa-tion of such an architecture. Operational and techno-logical risks remain as well, both of which are intendedto be ”bought down” in the frame of the demonstra-tor mission. A careful phasing of demonstrator andtarget architectures remain a key programmatic chal-lenge for the participating agencies in order to proposean attractive program to their respective stakeholders.

References

[1] International Space Exploration CoordinationGroup, “The Global Exploration Roadmap,” Au-gust 2013.

[2] Gerstenmaier, William H. “Progress in Definingthe Deep Space Gateway and Transport Plan”,Presentation to the NASA Advisory Council,March 2017.

[3] Whitley, R. and Martinez, R. “Options for StagingOrbits in Cislunar Space.” 2015 IEEE AerospaceConference, 2015.

[4] National Research Council, “The Scientific Con-text for Exploration of the Moon: Final Report”,ISBN: 0-309-10920-5, 2007

[5] Lunar Exploration Analysis Group, “The LunarExploration Roadmap: Exploring the Moon in the21st Century: Themes, Goals, Objectives, Investi-gations, and Priorities”, 2016

[6] O. M. Kamps, E. J. Allender, N. V. Almeida, J.Cook, J. J. Ende, S. Mazrouei, C. Orgel, T. Slezak,A. J. Soini, and D. A. Kring (2017), “Explorationof South Polar Region of the Moon: Tele-operatedTraverses”, Lunar and Planetary Science XLVIII,Abstract 1909.

[7] E.S.. Steenstra, D.J.P. Martin, F.E. McDonald,S.Paisamsombat, C. Venturino, S. OHara, A.Calzada-Diaz, S. Bottoms, M. K. Leader, K.K.Klaus, W. van Westrenen, D. H. Needham, D.A.

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Kring, “Analyses of robotic traverses and samplesites in the Schdinger basin for the HERACLEShuman-assisted sample return mission concept”,Advances in Space Research, Vol. 58, Iss. 6, pp.1050-1065, 2016.

[8] M. Landgraf, H. Ueno, O. A. Saprykin, “ExploitingHuman-Robotic Partnership Missions to Obtain aPristine Sample from Previously Inaccessible Siteson the Moon”, 1313, LPSC, 2015.

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