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Acta Astronautica

Acta Astronautica 68 (2011) 1201–1207

0094-57

doi:10.1

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journal homepage: www.elsevier.com/locate/actaastro

Lunette: A network of lunar landers for in-situ geophysical science$

John Elliott n, Leon Alkalai

Jet Propulsion Laboratory/California Institute of Technology, Pasadena, CA, United States

a r t i c l e i n f o

Article history:

Received 31 January 2010

Received in revised form

14 October 2010

Accepted 23 October 2010Available online 18 November 2010

Keywords:

Lunar lander

Network

65/$ - see front matter & 2010 Published by

016/j.actaastro.2010.10.024

s paper was presented during the 60th IAC in

esponding author. Tel.: +1 818 393 5992.

ail address: jelliott@jpl.nasa.gov (J. Elliott).

a b s t r a c t

Over the last 3 years, a team at JPL has worked to design a new concept for a small, low cost

lander applicable to a variety of in-situ lunar exploration activities. This concept, named

Lunette, originated as a design which would exploit potential excess capacity of EELV

launches by being compatible with the EELV Secondary Payload Adapter (ESPA). The original

Lunette mission concept would have allowed up to six low cost landers to be delivered to a

targeted region of the moon, with landings separated by a few km, allowing establishment of

a regional network with a single, shared launch. The original concept faced limits in the

extent of regional distribution of landing sites since all six landers were dependent on a single

solid rocket braking motor. In the last year the Lunette team has focused on a modification of

the original ESPA-based concept to a design that would allow launch of multiple individual

landers (each with its own braking stage) on a single launch vehicle, where each lander would

be capable of independent targeting and landing. With such an implementation, the entire

lunar surface could be accessed for establishment of network nodes that could enable high

priority geophysical measurements on a scale not seen since Apollo. The present paper

discusses the current state of the design of the Lunette geophysical network lander, as well as

describing mission design, science operations, and an innovative design solution allowing the

lander to take critical data continuously, even over the lunar night, without the need for

radioisotope power systems.

& 2010 Published by Elsevier Ltd.

1. Introduction

The original concept for the Lunette lander began as anexercise to develop a low cost lunar mission that could becarried as an Evolved Expendable Launch Vehicle (EELV)secondary payload. Emphasis was placed on simplicity ofdesign and operations and maximization of versatility toperform a wide variety of missions [1]. The design of theoriginal Lunette lander focused on a lander vehicle thatcould be mounted on the EELV secondary payload adapter(ESPA), allowing up to six landers to be launched as a singlesecondary payload on a moon-bound EELV. The six landerswere to fly to the moon as an integrated unit with the ESPAring, which itself housed a large solid rocket motor (SRM) to

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perform a braking burn over the lunar surface. At termina-tion of the SRM burn, the landers would separate from theESPA and each would then activate its own liquid propul-sion system to perform a targeted descent and landing,resulting in a network of up to six landers spread over anarea of 10–20 km on the lunar surface. This baselinemission was initially devised as a ‘‘site survey’’ to precedeestablishment of a permanent human outpost at the lunarSouth Pole, on the rim of Shackleton crater (Fig. 1).

During the development of this initial Lunette missionconcept, the team realized that the limitations of the ESPA-based implementation would limit applicability of the con-cept to relatively local ‘‘regional’’ network science. It wasapparent that more global network missions would requireindividually targeted landers, capable of being sent todestinations spread over the entire surface of the moon.Initially the team conceived of a simple adaptation of theESPA-based design which retained the lander configurationdevised to fit the ESPA envelope, but now mated to an

Fig. 1. Conceptual site survey mission overview.

Fig. 2. Lunette design evolution.

J. Elliott, L. Alkalai / Acta Astronautica 68 (2011) 1201–12071202

individual SRM for braking as demonstrated on the Surveyorprogram of the 1960s [2]. The small size of the lander/SRMstack and low mass of each system still allowed multiplelanders to be delivered with a single, now dedicated, EELVlaunch. In this design variation, each lander would separatefrom the launch vehicle following trans-lunar injection andproceed to the moon as an individual flight system. Eachlander could be targeted for a landing site anywhere on thelunar surface. Deployable solar arrays would be incorporatedto accommodate operations at any latitude.

As interest began to grow in establishment of geophy-sical networks for lunar science, the team investigatedapplicability of the individual lander design to an earlyconcept being considered by NASA for the InternationalLunar Network (ILN) project. This model mission consistedof two landers: one at the lunar South Pole and one at thelunar North Pole. A mission design and model payload suitewas developed and an estimate of mission cost determinedthrough JPL’s Team X. During the course of this study itbecame evident that although the performance of the

lander design was adequate to the mission needs, theconfiguration, tightly constrained as it was by the limita-tions of the ESPA payload envelope, was not ideal for theincorporation and deployment of the desired instrumentsuite. Additionally, the original lander design, while cap-able of surviving over the extended lunar night andoperating for years on the surface, was not designed totake data during night-time periods. Solutions to theseshortcomings resulted in a redesigned lander, with sim-plifications such as fixed legs, and internal augmentation ofthe Command and Data Handling (CADH) and Powersubsystems to provide a new design that could operatecontinuously, taking and accurately time-tagging sciencedata even over the lunar night. In addition, the thermalcontrol subsystem was refined to enable the elimination ofall radioisotope heater units (RHUs) which had previouslybeen relied upon for overnight survival heating. An over-view of the evolution of the Lunette design, including aninvestigation of an optional controllable SRM-based land-ing system, is shown in Fig. 2. The present paper discusses

J. Elliott, L. Alkalai / Acta Astronautica 68 (2011) 1201–1207 1203

the current state of the Lunette design, in the context of aglobal lunar geophysical network mission.

2. Mission concept

The Lunette geophysical mission study was basedclosely on the results of work performed by the ILN scienceDefinition Team (SDT) [3]. The Lunette team took the ILNSDT final report as a model for the mission concept,incorporating all of the science objectives and investiga-tions. The major divergence is in the desire for landing siteson the lunar far-side, which was considered beyond thescope of the JPL study because of the absence of telecomrelay assets. A mission concept was developed that couldaccommodate three near-side landing sites, chosen basedon criteria laid out in the ILN SDT final report.

The Lunette mission would begin with the launch ofthree lander/SRM flight systems on a single EELV. The stackwould be placed on a lunar trajectory by the launch vehicle,following which the individual flight systems would sepa-rate from the upper stage and perform a low energytransfer to the moon. The flight systems would arrive atLunar Lagrange point 2 (LL2) approximately 3 months afterlaunch, entering a Lissajous orbit around LL2. The landerswould be dispatched to the moon from this orbit over thecourse of several weeks, allowing each lander to land in themorning at its targeted site and complete its deploymentsand initial checkout operations before the next landerarrives. On lunar approach, the SRM would fire to slowthe lander to about 100 m/s horizontal velocity at analtitude of about 2 km above the lunar surface. The SRMwould then be jettisoned and the lander liquid propulsionsystem would be activated to perform final descent andlanding at the targeted site. Precision landing is notrequired for this mission, but the landers would beequipped with hazard detection and avoidance capability,based on a landing radar and terrain-following cameras.

Landing would be followed by deployment of theinstrument suite. The complement of instruments includesall of the investigations recommended by the ILN SDT and ismade up of a seismometer, a mole capable of placing a heatflow probe to a depth of 3 m, an electromagnetic instru-ment suite (comprised of two magnetometers, two elec-trometers, and a Langmuir probe), and a retroreflector. Ofthese, the seismometer, the mole and the retroreflector aredeployed on the lunar surface at a distance from the landerby a robotic arm. The electromagnetic suite deploys frombooms mounted to the lander body. All deployments wouldbe completed on the first lunar day for each lander, andscience data taking would commence following checkout.During daylight hours all lander systems would be active,with science data downlink direct-to-Earth through thelander’s S-band communication subsystem.

At sunset the lander would enter its power conservingnight-time mode. In this mode only the instruments and anevent timer module (ETM) remain powered, along withminimal electric heaters to maintain electronics abovesurvival temperatures (420 1C). The ETM, unique to thisdesign, serves as an interface between the instruments andthe CADH subsystem. Data collected by the instruments isdumped to the ETM for 1 min out of every hour, where it is

stored in non-volatile memory and time-tagged by theETM’s highly stable chip-scale atomic clock. At dawn, whensolar power is once again available, the remaining landersubsystems would be activated, time-tagged science datawould be passed to the CADH, and returned to Earththrough the telecommunications link. Operating in thismanner the landers are designed to continue their geo-physical monitoring mission for several years on the lunarsurface.

3. Conceptual mission design

The driving consideration in designing the Lunettetrajectory was the desire to land three landers at threegeographically distributed landing sites on the near side ofthe moon. Further, it was desired to ensure that each landerwould land at local morning, and that landings could beseparated in time sufficient to allow all critical activities tobe completed on one lander before initiating the nextlanding operation. The trajectory type chosen for transferof the Lunette flight systems to the moon is a low energytransfer, similar to that being planned for the upcomingGRAIL mission [4]. This option makes use of a longer flighttime to allow third-body perturbations from the Sun todecrease the arrival velocity at the moon. It also has theadvantage of very flexible launch planning, providing �21launch opportunities with a �24 day launch period, whichwould repeat every month with minor variations. For thebaseline mission the three flight systems would be launchedon an EELV to a C3 of about -0.5 km2/s2. The flight systemswould separate from the launch vehicle soon after launch,beginning a low energy transfer from the Earth to the moonvia the Sun–Earth L1 (EL1) vicinity (Fig. 3). This trajectoryhas the additional benefit of avoiding eclipses during cruise.Trans-lunar cruise duration would be between 80 and 100days, depending on launch date within the launch period.Cruise would include two deterministic maneuvers perflight system, with an additional three planned statisticalmaneuvers. Total deterministic and statistical DV for thecruise has been budgeted at �50 m/s for each flight system,based on the similar trajectories developed for GRAIL [5]. Tosimplify operations, it is planned that no two maneuversamong all three flight systems will be executed within a dayof each other.

Rather than landing directly on the moon, the threeflight systems will enter separate Lissajous staging orbitsaround LL2, each of which is designed to accommodate thatlander’s destination. Insertion into these orbits requires aminimal (o1 m/s) insertion burn. The staging orbit willallow the landers to be deployed individually to theirlanding sites, ensuring optimal landing times and allowingcritical surface operations to be completed before the nextlander arrives.

The orbits have been designed such that no determi-nistic maneuvers are required to depart and begin thedescent. An example descent scenario for each of threelanders is shown in Fig. 4. In these scenarios Lander 1 wouldperform a direct descent from LL2 to its landing site.Descent duration would be �7 days. Lander 2 wouldperform a 21 day descent from LL2 through three loopsaround the moon. Each phasing loop would shift the

Fig. 3. Example trans-lunar trajectory.

Fig. 4. Example lander descent trajectories.

J. Elliott, L. Alkalai / Acta Astronautica 68 (2011) 1201–12071204

lander’s trajectory until it is in a position to land at Site 2.Lander 3’s descent would also be a 21 day trajectory, verysimilar to Lander 2.

Each lander would approach its landing site at a shallowpath angle of �5 to �101. The braking burn would beinitiated at �13 km altitude and about 70 km from thelanding site.

4. Flight system design concept

As with the original Lunette concept, the currentflight system design is aimed at simplicity, low cost, and

low risk. The majority of the flight system is made up offlight proven off-the-shelf or catalog subsystems andcomponents.

An overview of the flight system in its cruise config-uration is shown in Fig. 5. Given the elimination of theconstraints imposed by the earlier Lunette accommodationon the ESPA, a simplification has been introduced byeliminating deployable landing legs in favor of six fixedlegs. The figure shows the lander attached to its Star 27SRM. The Star 27 is attached to the lander body by aseparation system. A star tracker is also attached to thestructural support for the SRM.

Solid Rocket Motor

Solar Arrays (6)

DeploymentArm

Thruster Cluster (3)

Star tracker

Radiator with Reflector

Fig. 5. Conceptual flight system in cruise configuration.

Fig. 6. Launch configuration shown in representative atlas V 5-m fairing.

J. Elliott, L. Alkalai / Acta Astronautica 68 (2011) 1201–1207 1205

The three flight systems would be supported duringlaunch by a custom launch vehicle adapter (LVA). Theconceptual launch configuration as shown in Fig. 6 leavessubstantial open volume in an EELV 5-m fairing. Separation

from the LVA would be accomplished through the use ofcommercially available motorized Lightband separationsystems.

The lander propulsion system retains the Phoenixheritage monopropellant landing thrusters and pulse-modulated throttling design, with two 267 N landingthrusters mounted on each of three legs. In addition, acluster of three reaction control system (RCS) 0.9 N thrus-ters are co-located with the landing thrusters on thesethree legs. The body of the spacecraft houses the hydrazinefuel tank, sized to hold 45 kg of fuel for the mission. Allavionics are co-located in a warm electronics box (WEB)enclosure at the top of the lander body. These avionics aremated to a radiator plate with louvers for thermal control. Aparabolic reflector structure mounted over this radiatorserves to protect the radiator from direct sunlight over thecourse of the lunar day. At night, a motorized thermal coveris provided to enclose the radiator and louvers to minimizeheat losses. Thermal design of the WEB has been opti-mized to limit heat loss to no more than 3 W during nightoperations using multilayer insulation and low conductiv-ity materials to isolate the WEB from the lander structure.In combination with the low power nighttime operationsscenario enabled by the ETM, this implementation shouldallow the lander to collect data over the full course of thelunar night (up to 16 earth days) using a total of �36 kg(including contingency) of Li-ion batteries with less than60% depth of discharge.

Solar arrays are provided on the six sides of the landerbody. These solar arrays are hinged at the top and would bedeployed after landing to an angle optimum to their lunarlatitude. This feature retains the flexibility of the design toallow operation from the poles (no need for deployment) tothe equator (deployment to 901). Angle of deploymentwould be fixed for each lander, set prior to launch accordingto the lander’s lunar destination.

Instruments to be deployed following landing would bemounted on the lander’s legs during cruise. Three of theinstruments, the seismometer, the mole/heat flow probe,and the retroreflector, require deployment away from thelander for optimum performance. These three instrumentswould be deployed remotely by a robotic arm mounted tothe leg structure. This arm, equipped with a simple forkedend effecter, would engage with a deployment pin on thetop of each instrument. The instrument would be lifted, andplaced in a suitable spot approximately 2 m from the landerbody. The arm would be used to deploy these instrumentsthe first day after landing, and would not be required for theremainder of the mission. The remaining instruments ofthe electromagnetic suite are mounted on the lander body.Two magnetometers and an electrometer would bedeployed on a boom deployed horizontally, and an addi-tional magnetometer would deploy on a second boomorthogonal to the first. Finally, a Langmuir probe is hard-mounted on a short mast on the top deck of the lander. Thedeployed configuration of the lander and instruments isshown in Fig. 7.

The baseline flight system master equipment list (MEL)is shown in Table 1. Masses were estimated by eachsubsystem based on design and component specificationswith each subsystem carrying a contingency reflecting

Fig. 7. Conceptual lander configuration showing instrument suite.

Table 1Flight system MEL.

CBE (kg) Cont. (%) Total (kg)

LanderInstruments 12.6 27% 16.0

C&DH 7.4 16% 8.6

Power 34.6 18% 40.8

Telecom 3.9 15% 4.5

Structures 39.9 30% 51.9

Thermal 6.6 27% 8.3

Propulsion 32.0 33% 42.4

GN&C 2.8 20% 3.3

Lander total 139.8 26% 175.8System Margin 24.0

Dry mass total 43% 199.8Propellant 45.0

Wet mass total 244.8

SRM/CarrierStructures 12.8 30% 16.6

Thermal 2.1 30% 2.7

Propulsion 29.5 10% 32.5

GN&C 2.5 10% 2.8

Carrier total 46.9 16% 54.512.5

Dry mass total 43% 67.0Propellant 338.0

Wet mass total 405.0

Single flight systemLander 244.8

Carrier (wet) 405.0

Wet mass total 649.9

J. Elliott, L. Alkalai / Acta Astronautica 68 (2011) 1201–12071206

design maturity in accordance with JPL Flight ProjectPractices. In addition to subsystem contingency, a ‘‘systemmass margin’’ is added to each flight system element tobring the total dry mass contingency to 43% (30% marginper JPL Design Principles).

The Carrier stage portion of the MEL tallies the masses ofthe Star 27 SRM and its attachment structure and the startracker. 338 kg of solid propellant is carried in the Star 27 to

provide a braking DV capability of �2300 m/s to theintegrated flight system. Total launch mass of eachlander/SRM stack is estimated at about 650 kg. The totalflight system launch mass for three flight systems would be1950 kg, making a three lander mission compatible witheven the smallest EELV. The lander dry mass with margin is200 kg, of which 16 kg constitutes the science payload.

5. Summary and conclusions

The Lunette study began as an exercise to develop a lowcost lander implementation that could be launched as anEELV secondary payload. The original design succeeded atfleshing out that concept, with development of a capablelow risk architecture that could enable a number ofregional network science missions at a minimal cost.Evolution of the concept to apply to wider mission applic-ability has led the design team to a more versatileimplementation. The current Lunette flight system conceptretains the simplicity of the original, but provides greatlyincreased science capability by allowing globally distrib-uted networks to be established with a single launch. Themission described in this paper takes as its inspiration theInternational Lunar Network concept currently beingexplored by NASA; however, the application of this landerconcept could enable a wide variety of landed missions. Thecombination of simplicity, low mass, small size, and globalaccess makes a powerful case for expansion of lunar surfacescience in the next decade.

Acknowledgements

The authors wish to thank and acknowledge the Lunettedesign team, including Melissa Jones, Tim McElrath, BobMiyake, Paul Timmerman, Steve Kondos, Matt Spaulding,Vince Randolph, Dwight Geer, Joseph Smith, Mike Gallagher.

All of the research described in this paper was carriedout at the Jet Propulsion Laboratory, California Institute of

J. Elliott, L. Alkalai / Acta Astronautica 68 (2011) 1201–1207 1207

Technology, under a contract with the National Aeronau-tics and Space Administration.

References

[1] J.O. Elliott, L. Alkalai, Lunette: a low-cost concept enabling multi-lander lunar science and exploration missions, in: Proceedings of the

59th International Astronautics Congress, Glasgow, Scotland, 2008.[2] Surveyor Program Results, NASA SP-184, National Aeronautics and

Space Administration, Washington, DC, 1969.

[3] Science Definition Team for the ILN Anchor Nodes, ILN Final Report,NASA, 2009, /http://lunarscience.arc.nasa.gov/articles/international-lunar-network-issues-2009-final-reportS.

[4] R.B. Roncoli, K.K. Fujii, Mission design overview for the gravityrecovery and interior laboratory (GRAIL) mission, in: Proceedings ofthe AIAA Guidance, Navigation, and Control Conference, No. AIAA2010-8383, AIAA, Toronto, Ontario, Canada, August 2–5, 2010.

[5] M.J. Chung, S.J. Hatch, J.A. Kangas, S.M. Long, R.B. Roncoli, T.H. Sweetser,Trans-lunar cruise trajectory design of GRAIL (gravity recovery andinterior laboratory) mission, in: Proceedings of the AIAA Guidance,Navigation, and Control Conference, No. AIAA 2010-8384, AIAA, Toronto,Ontario, Canada, August 2–5, 2010.