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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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17 records
All sourcesNASA NTRSDOE OSTI
DOE OSTI2024

Radioisotope Power System Cask Trailer Chiller Low Heat Load Study

This report documents the study of the Radioisotope Thermoelectric Generator Transport System (RTGTS) chillers capability to transport a Radioisotope Power System (RPS) with low heat loads in the cask. Tests of the RTGTS chillers included setting the glycol/water (coolant) temperature to a control set point of 40° and 60°F while applying heat loads ranging from 2500 to 0 W heat load. The tests performed showed that the chiller systems could handle the entire range of heat loads. However, a noise/vibration started to occur when the set point was 60°F with low heat loads. Three recommended changes to the RTGTS chillers are recommended with varying implementation windows. The first is an operational change that can be implemented immediately. The RTGTS chillers should have a set point of 40°F when transporting an RPS less than 1500 W. The second is a short-term change (within the next two years) augmenting RPS heat by utilizing one RTGTS onboard heater (2500 W) while transporting an RPS with less than 1500 W which will also require a change to the Instrument and Data Acquisition System (IDAS) to allow monitoring and possible control of power to the onboard heaters. However, if the heater does fail during transport, no emergency actions or immediate repairs will be required due to the chillers’ ability to handle low heat loads. Running a 2500 W heater when transporting low heat loads aids in prolonging the life of the chillers. The remaining recommended long-term change is to replace the current chillers with a new portable or permanently installed chiller system in the next five to ten years. This time frame takes advantage of the proven reliability and durability of the current chillers while the new chillers are phased in and proven to the RPS program.

42 ENGINEERING↗
NASA NTRS1990

Accommodation of Nuclear Power and Propulsion Concepts

The use of nuclear systems for propulsion and power are being examined as system options for implementing the lunar and Mars human exploration missions currently being studied by NASA. Systems might include nuclear electric propulsion (NEP) and nuclear thermal rocket (NTR) vehicles, operating reactors on coorbiting platforms, radioisotope thermoelectric generators, and others. The space station, as a transportation node, would have to store, assemble, launch and refurbish elements containing these systems. Care must be taken to safeguard humans from the radiation imposed by these systems, in addition to the naturally occuring background of the space environment. Key issues need to be identified early to enable their proper consideration in planning activities and the baseline space station design. A study was conducted over the past year with Texas A&M University to identify and explore key issues and quantify findings in a way useful to the Space Station Program.

Stevenson, Steven M.↗
NASA NTRS1989

Summary of miscellaneous hazard environments for hypothetical Space Shuttle and Titan IV launch abort accidents

The various analyses described here were aimed at obtaining a more comprehensive understanding and definition of the environments in the vicinity of the Radioisotope Thermal Generator (RTG) during certain Space Transportation System (STS) and Titan IV launch abort accidents. Addressed here are a number of issues covering explosion environments and General Purpose Heat Source Radioisotope Thermoelectric Generator (GPHS-RTG) responses to those environments.

Eck, M.↗
NASA NTRS1989

Shuttle data book: SRM fragment velocity model. Presented to the SRB Fragment Model Review Panel

This study was undertaken to determine the velocity of fragments generated by the range safety destruction (RSD) or random failure of a Space Transportation System (STS) Solid Rocket Motor (SRM). The specific requirement was to provide a fragment model for use in those Galileo and Ulysses RTG safety analyses concerned with possible fragment impact on the spacecraft radioisotope thermoelectric generators (RTGS). Good agreement was obtained between predictions and observations for fragment velocity, velocity distributions, azimuths, and rotation rates. Based on this agreement with the entire data base, the model was used to predict the probable fragment environments which would occur in the event of an STS-SRM RSD or randon failure at 10, 74, 84 and 110 seconds. The results of these predictions are the basis of the fragment environments presented in the Shuttle Data Book (NSTS-08116). The information presented here is in viewgraph form.

Source record↗
DOE OSTI2021

Modeling of Am-241 as an alternative fuel source in a radioisotope thermoelectric generator

Radioisotope thermoelectric generators (RTGs) have been utilized in the USA to power satellites and space exploration equipment for over half a century. In our preceding paper, Dustin and Borrelli (2021), we assessed the performance of 241 Am as an alternative to 238 Pu by modeling the specific decay-heat and gamma spectra of several potential radioisotopes in SCALE-ORIGEN, and comparing these with 238 Pu An analysis of heat output, heat stability, safety, availability and required mass showed that of the isotopes reviewed, 241 Am had the greatest potential for replacing 238 Pu, with additional design considerations. In our current work, we have assessed radiation transport for two configurations of an 241 Am powered Multi-Mission Radioisotope Thermoelectric Generator (MMRTG), due to the potential health effects for operators, fabricators, and others with the potential to be exposed to neutrons and lower energy gamma rays. We have modeled the use of americium as a fuel source replacement for plutonium in current MMRTG designs. The assessment has been performed using MCNP6, in terms of neutron and gamma emissions that might interfere with sensitive equipment and power system production. The radiation transport behavior in terms of dose rate of the 241 Am fueled models is compared with that of a 238 Pu fueled model. Finally, additional work was performed to validate modeling simplifications with regards to thermocouple material configurations.

11 NUCLEAR FUEL CYCLE AND FUEL MATERIALS↗
NASA NTRS1979

Radioisotope thermoelectric generator cooling in the Shuttle bay

The paper describes a Shuttle-integrated radioisotope thermoelectric generator (RTG) that consists primarily of a pump package and plumbing connected directly to the Shuttle payload heat exchanger. The RTG utilizes on-board water evaporative cooling capability, which is normally used for ascent, entry, and for supplementing the radiators. Attention is given to the RTG cooling concepts which include: (1) an active thermal cooling system (ATCS), where two Freon-21 loops operate simultaneously to transport heat from the Orbiter subsystem and payloads through liquid-to-liquid heat exchangers and pin-fin coldplates to four heat sinks, and (2) an atmosphere revitalization system (ARS) which provides for thermal, pressure, and contaminate control of the crew cabin and its equipment. The use of a payload heat exchanger to reduce weight, cost and complexity associated with an independent cooling system was investigated in detail.

Stimpson, L. D.↗
NASA NTRS1992

Toward the last frontier - A strategy for the evolutionary development of space nuclear power systems

A number of exciting mission opportunities are being considered for the 21st century, including advanced robotic science missions to the outer planets and beyond, human exploration of the Moon and Mars, and advanced space transportation systems. All of these missions will require some form of nuclear power; however, it is clear that current budgetary constraints preclude developing many different types of space nuclear power systems. This paper reviews the specific civil space missions which have been identified, the power levels and lifetimes required, and the technologies available. From this an evolutionary space nuclear power program is developed which builds upon the experience of radioisotope thermoelectric generators, dynamic isotope power systems, and space nuclear reactors. It is strongly suggested that not only does this approach make technical and budgetary sense but that it is consistent with the normal development of new technologies.

Gary L Bennett↗
NASA NTRS2022

COMPASS Final Report: 2008 International Lunar Network (ILN)

The Compass team developed three designs for a soft lander for the International Lunar Network (ILN) using different power systems. These systems included solar arrays and batteries and two radioisotope power system (RPS) options scaled down from previous designs. These are a half Stirling radioisotope generator and a one-eighth Radioisotope Thermoelectric Generator (RTG). The COMPASS design consists of a soft-lander capable of network science (seismic, magnetospheric, laser reflector, and thermal transfer) that can land anywhere on the near side of the Moon and operate for at least six years. The baseline version, using only batteries and solar arrays, requires a very large battery (~80 kg) for the roughly 20 W night-time power requirements. The lander launch mass (with solid rocket breaking motor) is almost 900 kg. With such a large battery, the lander will require a launcher larger than the desired Minotaur V; a Taurus II (being developed for NASA Commercial Orbital Transportation Services (COTS)) or larger (Delta II or Atlas V) would be needed. The two RPS cases, half an Advanced Stirling Radioisotope Generators (ASRG) and one-eighth RTG have much smaller launch masses of 500 and 600 kg respectively. (The one-eighth RTG case lacks the power for the full science package, so a solar array battery is still needed.) Unfortunately, the Minotaur V’s launch capability is still not large enough (~430 kg to trans-lunar injection (TLI)) for the RPS systems. Consequently, a fourth case was developed to fly a limited science floor suite of instruments (seismometer only) which reduced the RPS Lander launch mass to fit on a Minotaur V.

Radioisotope Power Systems↗
NASA NTRS2012

Multi-Mission Radioisotope Thermoelectric Generator Heat Exchangers for the Mars Science Laboratory Rover

The addition of the Multi-Mission Radioisotope Thermoelectric Generator (MMRTG) to the Mars Science Laboratory (MSL) Rover requires an advanced thermal control system that is able to both recover and reject the waste heat from the MMRTG as needed in order to maintain the onboard electronics at benign temperatures despite the extreme and widely varying environmental conditions experienced both on the way to Mars and on the Martian surface. Based on the previously successful Mars landed mission thermal control schemes, a mechanically pumped fluid loop (MPFL) architecture was selected as the most robust and efficient means for meeting the MSL thermal requirements. The MSL heat recovery and rejection system (HRS) is comprised of two Freon (CFC-11) MPFLs that interact closely with one another to provide comprehensive thermal management throughout all mission phases. The first loop, called the Rover HRS (RHRS), consists of a set of pumps, thermal control valves, and heat exchangers (HXs) that enables the transport of heat from the MMRTG to the rover electronics during cold conditions or from the electronics straight to the environment for immediate heat rejection during warm conditions. The second loop, called the Cruise HRS (CHRS), is thermally coupled to the RHRS during the cruise to Mars, and provides a means for dissipating the waste heat more directly from the MMRTG as well as from both the cruise stage and rover avionics by promoting circulation to the cruise stage radiators. A multifunctional structure was developed that is capable of both collecting waste heat from the MMRTG and rejecting the waste heat to the surrounding environment. It consists of a pair of honeycomb core sandwich panels with HRS tubes bonded to both sides. Two similar HX assemblies were designed to surround the MMRTG on the aft end of the rover. Heat acquisition is accomplished on the interior (MMRTG facing) surface of each HX while heat rejection is accomplished on the exterior surface of each HX. Since these two surfaces need to be at very different temperatures in order for the fluid loops to perform efficiently, they need to be thermally isolated from one another. The HXs were therefore designed for high in-plane thermal conductivity and extremely low through-thickness thermal conductivity by using aluminum facesheets and aerogel as insulation inside a composite honeycomb core. Complex assemblies of hand-welded and uniquely bent aluminum tubes are bonded onto each side of the HX panels, and are specifically designed to be easily mated and demated to the rest of the RHRS in order to ease the integration effort.

Mastropietro, A. J.↗
NASA NTRS2020

Innovative Thermal Management & Control to Surmount Challenges of Exploring Ocean Worlds on Europa and Enceladus

Probes to penetrate the thick ice shells of our solar system’s Ocean Worlds have been studied for nearly 20 years, since scientific evidence strongly suggested a subsurface ocean on the Jupiter moon called Europa. There is keen scientific interest in exploring four significant themes on such proposed missions: 1) Geodynamics, 2) Geochemistry, 3) Habitability, and 4) Life Detection. The ice shells of Ocean Worlds are predicted to be up to 40 km thick; they exhibit extreme thermal environments, with ice temperatures from 100 K to 270 K, and extreme pressure environments from vacuum to 53 MPa. Jet Propulsion Laboratory has conducted a broad-look investigation of proposed mission concepts to Europa to identify the significant technology and operational challenges of Europa ice-penetration. The thermal-mechanical system (TMS) of an ice penetration probe (IPP) designed to access the ocean of an icy moon using radioisotope thermoelectric generators for heat and power faces technological hurdles exacerbated by severe thermal and volume constraints. This study identified thermal management and control (TMC) challenges that are strongly linked to: ice penetration start-up, mobility and navigation in the ice, communications while in the ice sheet, and detecting and avoiding in-ice hazards. The major objectives of the TMC system are: 1) Absorb internal thermal energy from the IPP radioisotope power source, 2) Maintain liquid water conditions around the IPP at all times, 3) Manage and control thermal flows from probe nose to tail, and 4) Provide pressure containment for all internal probe components. This work discusses the baseline TMC system architecture and design developed to accomplish these objectives, and survive and transit the extreme ice thicknesses in pursuit of Icy/Ocean Worlds science goals. The proposed TMC system consisting of an internal pumped two-phase fluid loop “thermal bus” for thermal energy capture, variable conductance heat pipe system for passively adaptive thermal energy transport around the probe, and water jetting system for ice cutting is described and discussed. Critical testing performed to date is described.

Lee, Kuan-Lin↗
NASA NTRS1992

Pressurized Lunar Rover (PLR)

The objective of this project was to design a manned pressurized lunar rover (PLR) for long-range transportation and for exploration of the lunar surface. The vehicle must be capable of operating on a 14-day mission, traveling within a radius of 500 km during a lunar day or within a 50-km radius during a lunar night. The vehicle must accommodate a nominal crew of four, support two 28-hour EVA's, and in case of emergency, support a crew of six when near the lunar base. A nominal speed of ten km/hr and capability of towing a trailer with a mass of two mt are required. Two preliminary designs have been developed by two independent student teams. The PLR 1 design proposes a seven meter long cylindrical main vehicle and a trailer which houses the power and heat rejection systems. The main vehicle carries the astronauts, life support systems, navigation and communication systems, lighting, robotic arms, tools, and equipment for exploratory experiments. The rover uses a simple mobility system with six wheels on the main vehicle and two on the trailer. The nonpressurized trailer contains a modular radioisotope thermoelectric generator (RTG) supplying 6.5 kW continuous power. A secondary energy storage for short-term peak power needs is provided by a bank of lithium-sulfur dioxide batteries. The life support system is partly a regenerative system with air and hygiene water being recycled. A layer of water inside the composite shell surrounds the command center allowing the center to be used as a safe haven during solar flares. The PLR 1 has a total mass of 6197 kg. It has a top speed of 18 km/hr and is capable of towing three metric tons, in addition to the RTG trailer. The PLR 2 configuration consists of two four-meter diameter, cylindrical hulls which are passively connected by a flexible passageway, resulting in the overall vehicle length of 11 m. The vehicle is driven by eight independently suspended wheels. The dual-cylinder concept allows articulated as well as double Ackermann steering. The primary power of 8 kW is supplied by a dynamic isotope system using a closed Brayton cycle with a xenon-hydrogen mixture as the working fluid. A sodium-sulfur battery serves as the secondary power source. Excess heat produced by the primary power system and other rover systems is rejected by radiators located on the top of the rear cylinder. The total mass of the PLR 2 is 7015 kg. Simplicity and low total weight have been the driving principles behind the design of PLR 1. The overall configuration consists of a 7-m-long, 3-m-diameter cylindrical main vehicle and a two-wheeled trailer. The cylinder of the main body is capped by eight-section, faceted, semi-hemispherical ends. The trailer contains the RTG power source and is not pressurized. The shell of the main body is constructed of a layered carbon fiber/foam/Kevlar sandwich structure. Included in the shell is a layer of water for radiation protection. The layer of water extends from the front of the rover over the crew compartment and creates a safe haven for the crew during a solar flare-up. The carbon fiber provides the majority of the strength and stiffness and the Kevlar provides protection from micrometeoroids. The Kevlar is covered with a gold foil and multi-layer insulation (MLI) to reduce radiation degradation and heat transfer through the wall. A thin thermoplastic layer seals the fiber and provides additional strength.

Creel, Kenneth↗
NASA NTRS1990

Draft Environmental Impact Statement for the Ulysses Mission (Tier 2)

This Draft Environmental Impact Statement (DEIS) addresses the environmental impacts which may be caused by the preparation and operation of the Ulysses spacecraft, including its planned launch on the Space Transportation System (STS) Shuttle and the alternative of canceling further work on the mission. The launch configuration will use the STS/Inertial Upper Stage (IUS)/Payload Assist Module-Special(PAM-S) combination. The Tier 1 EIS included a delay alternative which considered the Titan 4 launch vehicle as an alternative booster stage for launch in 1991 or later. However, the U.S. Air Force, which procures the Titan 4 for NASA, could not provide a Titan 4 vehicle for the 1991 launch opportunity because of high priority Department of Defense requirements. The only expected environmental effects of the proposed action are associated with normal Shuttle launch operations. These impacts are limited largely to the near-field at the launch pad, except for temporary stratospheric ozone effects during launch and occasional sonic boom effects near the landing site. These effects have been judged insufficient to preclude Shuttle launches. In the event of (1) an accident during launch, or (2) reentry of the spacecraft from earth orbit, there are potential adverse health and environmental effects associated with the possible release of plutonium dioxide from the spacecraft's radioisotope thermoelectric generators (RTG). The potential effects considered in this EIS include risks of air and water quality impacts, local land area contamination, adverse health and safety impacts, the disturbance of biotic resources, impacts on wetland areas or areas containing historical sites, and socioeconomic impacts. Intensive analysis of the possible accidents associated with the proposed action are underway and preliminary results indicate small health or environmental risks. The results of a Final Safety Analysis Report will be available for inclusion into the final EIS.

Source record↗
NASA NTRS1988

Draft environmental impact statement for the Galileo Mission (Tier 2)

This Draft Environmental Impact Statement (DEIS) addresses the environmental impacts which may be caused by the preparation and operation of the Galileo spacecraft, including its planned launch on the Space Transportation System (STS) Shuttle and the alternative of canceling further work on the mission. The launch configuration will use the STS/Inertial Upper Stage (IUS)/Payload Assist Module-Special (PAM-S) combination. The Tier 1 EIS included a delay alternative which considered the Titan 4 launch vehicle as an alternative booster stage for launch in 1991 or later. However, the U.S. Air Force, which procures the Titan 4 for NASA, could not provide a Titan 4 vehicle for the 1991 launch opportunity because of high priority Department of Defense requirements. The only expected environmental effects of the proposed action are associated with normal Shuttle launch operations. These impacts are limited largely to the near-field at the launch pad, except for temporary stratospheric ozone effects during launch and occasional sonic boom effects near the landing site. These effects have been judged insufficient to preclude Shuttle launches. In the event of: (1) an accident during launch, or (2) reentry of the spacecraft from earth orbit, there are potential adverse health and environmental effects associated with the possible release of plutonium dioxide from the spacecraft's radioisotope thermoelectric generators (RTG).

Source record↗
NASA NTRS1994

Design of an unmanned Martian polar exploration system

The design of an unmanned Martian polar exploration system is presented. The system elements include subsystems for transportation of material from earth to Mars, study of the Martian north pole, power generation, and communications. Early next century, three Atlas 2AS launch vehicles will be used to insert three Earth-Mars transfer vehicles, or buses, into a low-energy transfer orbit. Capture at Mars will be accomplished by aerobraking into a circular orbit. Each bus contains four landers and a communications satellite. Six of the twelve total landers will be deployed at 60 deg intervals along 80 deg N, and the remaining six landers at 5 deg intervals along 30 deg E from 65 deg N to 90 deg N by a combination of retrorockets and parachutes. The three communications satellites will be deployed at altitudes of 500 km in circular polar orbits that are 120 deg out of phase. These placements maximize the polar coverage of the science and communications subsystems. Each lander contains scientific equipment, two microrovers, power supplies, communications equipment, and a science computer. The lander scientific equipment includes a microweather station, seismometer, thermal probe, x-ray spectrometer, camera, and sounding rockets. One rover, designed for short-range (less than 2 km) excursions from the lander, includes a mass spectrometer for mineral analysis, an auger/borescope system for depth profiling, a deployable thermal probe, and charge coupled device cameras for terrain visualization/navigation. The second rover, designed for longer-range (2-5 km) excursions from the lander, includes radar sounding/mapping equipment, a seismometer, and laser ranging devices. Power for all subsystems is supplied by a combination of solar cells, Ni-H batteries, and radioisotope thermoelectric generators. Communications are sequenced from rovers, sounding rockets, and remote sensors to the lander, then to the satellites, through the Deep Space Network to and from earth.

Baldwin, Curt↗
NASA NTRS2010

Lunar Surface Stirling Power Systems Using Isotope Heat Sources

For many years, NASA has used the decay of plutonium-238 (Pu-238) (in the form of the General Purpose Heat Source (GPHS)) as a heat source for Radioisotope Thermoelectric Generators (RTGs), which have provided electrical power for many NASA missions. While RTGs have an impressive reliability record for the missions in which they have been used, their relatively low thermal to electric conversion efficiency and the scarcity of plutonium-238 (Pu-238) has led NASA to consider other power conversion technologies. NASA is considering returning both robotic and human missions to the lunar surface and, because of the long lunar nights (14.75 Earth days), isotope power systems are an attractive candidate to generate electrical power. NASA is currently developing the Advanced Stirling Radioisotope Generator (ASRG) as a candidate higher efficiency power system that produces greater than 160 W with two GPHS modules at the beginning of life (BOL) (~32% efficiency). The ASRG uses the same Pu-238 GPHS modules, which are used in RTG, but by coupling them to a Stirling convertor provides a four-fold reduction in the number of GPHS modules. This study considers the use of americium-241 (Am-241) as a substitute for the Pu-238 in Stirling- convertor-based Radioisotope Power Systems (RPS) for power levels from tens of watts to 5 kWe. The Am-241 is used as a substitute for the Pu-238 in GPHS modules. Depending on power level, different Stirling heat input and removal systems are modeled. It was found that substituting Am-241 GPHS modules into the ASRG reduces power output by about one-fifth while maintaining approximately the same system mass. In order to obtain the nominal 160 W of electrical output of the Pu-238 ASRG requires 10 Am-241 GPHS modules. Higher power systems require changing from conductive coupling heat input and removal from the Stirling convertor to either pumped loops or heat pipes. Liquid metal pumped loops are considered as the primary heat transportation on the hot end and water pumped loop/heat pipe radiator is considered for the heat rejection side for power levels above 1 kWe.

Schmitz, Paul C.↗
NASA NTRS2009

Lunar Surface Stirling Power Systems Using Am-241

For many years NASA has used the decay of Pu-238 (in the form of the General Purpose Heat Source (GPHS)) as a heat source for Radioisotope Thermoelectric Generators (RTG), which have provided electrical power for many NASA missions. While RTG's have an impressive reliability record for the missions in which they have been used, their relatively low thermal to electric conversion efficiency (-5% efficiency) and the scarcity of Plutoinium-238 (Pu-238) has led NASA to consider other power conversion technologies. NASA is considering returning both robotic and human missions to the lunar surface and, because of the long lunar nights (14 earth days) isotope power systems are an attractive candidate to generate electrical power. NASA is currently developing the Advanced Stirling Radioisotope Generator (ASRG) as a candidate higher efficiency power system that produces greater than 160 watts with 2 GPHS modules at the beginning of life (BOL) (-30% efficiency). The ASRG uses the same Pu-238 GPHS modules, which are used in RTG, but by coupling them to a Stirling convertor provides a 4-fold reduction in the number of GPHS modules. This study considers the use of Americium 241 (Am-241) as a substitute for the Pu-238 in Stirling convertor based Radioisotope Power Systems (RPS) for power levels from 1 O's of watts to 5 kWe. The Am-241 is used as a replacement for the Pu-238 in GPHS modules. Depending on power level, different Stirling heat input and removal systems are modeled. It was found that substituting Am-241 GPHS modules into the ASRG reduces power output by about 1/5 while maintaining approximately the same system mass. In order to obtain the nominal 160 watts electrical output of the Pu-238 ASRG requires 10 Am-241 GPHS modules. Higher power systems require changing from conductive coupling heat input and removal from the Stirling convertor to either pumped loops or heat pipes. Liquid metal pumped loops are considered as the primary heat transportation on the hot end and water pumped loop/heat pipe radiator is considered for the heat rejection side for power levels above 1 kWe.

Schmitz, Paul C.↗
NASA NTRSDate not supplied

A Notional Artemis Lunar Surface Exploration Package (ArLSEP) based on the Gandalf Staff Platform

Introduction: The Artemis program is planning to deliver crew and cargo to the lunar surface, but there is no current package for supporting lunar in-struments and experiments similar to the Apollo Lunar Surface Exploration Package (ALSEP). This abstract provides a possible concept for such a package using the Gandalf Staff Platform as a common core. Gandalf Staff: The Gandalf Staff is an early prototype system developed over FY’21/FY’22 using NASA Science Technology Mission Directorate (STMD) Center Information Fund (CIF) grants to de-sign, build and test “proof-of-concept” components. These components include a 24v battery powered monopole that powers a suite of subsystems, including a Graphical User Interface (GUI) for crew, surface voice and data communications, Lunar Search and Rescue (LunaSAR) navigation and communications, LiDAR, field site external lighting, 360-degree camera, and a geothermal instrument for measuring sub-surface temperature gradient. The staff can be carried independently by an Extra-Vehicular Activity (EVA) astronaut, or can be mounted into a tripod for “hands free” support at a surface site being investigated. The staff can be attached to an external solar array and power storage system for long-duration operations. [1,2] ALSEP: An ASLEP flew on each mission Apollo 12 to Apollo 17. For Apollo 11, a simplified packaged called the Early Apollo Scientific Experiments Pack-age (EASEP) was flown. Each package included a “Central Station” that provided the power and communications connected to a variety of instruments and sensors. The power was provided by a Radioisotope Thermoelectric Generator (RTG) fueled by Plutoni-um-238 generating 70 watts of power (initially, decayed over time) [3]. The communications system provide for direct to Earth data transfer from the lunar surface. Each pack-age was stowed externally in the Lunar Module (LM) Scientific Equipment (SEQ) bay with a mass up to 163 kg (Apollo 17). The crew unloaded the ALSEP from the LM and deployed the instruments on the lunar surface. Although designed to operate for only 1 year, many sites operated for up to 8 years successfully [4]. The Active Seismic Experiment (ASE) included 3 geophones for detecting seismic waves created by mortars and thumpers deployed by the crew. Other active experiments measured the lunar atmosphere, the heat flow in the subsurface, the lunar gravity and potential gravity waves, the lunar magnetic field, the solar wind and plasma interactions in cislunar space. Passive experiments included collectors for dust and cosmic rays, and retroreflectors for precise measurements of distance using a laser from Earth. The ALSEP program continues to generate insights into lunar formation and evolution. ArLSEP Concepts: The lunar surface science package for the Artemis program will hopefully exceed the capability of the ALSEP. There are multiple issues for discussion leading to the design of a new ArLSEP, needing requirements definition from the science community, NASA mission architecture, and NASA budget planners. 1. Delivery Mechanism Two possible projects currently provide capability to deliver scientific cargo to the lunar surface: 1) the Commercial Lunar Payload Services (CLPS) [5] and the Human Landing System (HLS) [6, 7]. Each project is controlled by a different organization within NASA and budgeted with different criteria although both support lunar exploration. The HLS system delivers crew (and potentially cargo) to human landing sites. If an ArLSEP is “predeployed” to such a site, the design must include power (either from the vehicle or independently) to keep the electronics functioning until deployed by the crew. If an ArLSEP is delivered on a vehicle after the crew is present on the lunar surface, safety protocols require adequate distance from the humans for impact from descent propelled sur-face regolith ejecta. This distance can not exceed the capability of the crew to walk (if no rover) to the vehicle for ArLSEP deployment. 2. Overall Guidelines The general design of ArLSEP will likely follow the ALSEP with a common system for communications and power; however, significant architecture differences between Apollo and Artemis exist. Power: The RTG will not be available for early Artemis missions nor likely follow-on Lunar Exploration Transportation Services (LETS) missions [8]. Thus, ArLSEP power must be supplied by solar arrays with sufficient battery capability to “keep alive” necessary electronics during any lunar surface eclipse period. Communication: The Artemis program is developing a series of communications satellites for lunar orbit to provide surface transmission of data and voice to Earth. Called “LunaNET”, this network is component useful for ArLSEP since south polar locations may not always have direct “line-of-sight” to Earth [9]. 3. Concept of Operations (ConOps) The general ConOps for ArLSEP is to deliver the package to lunar surface before the crew arrives, and then have the crew deploy the package after some period of time. This requires coordinated design (for power systems) and launch window (for schedule) on both the cargo and crew missions. Once the ArLSEP is deployed, it will operate autonomously for a number of years. It should be designed to be EVA compatible for crew maintenance and upgrade. 4. Notional Design (for discussion purpose only) The landing site near the South Pole is expected to have no eclipse cycle exceeding 5 days, so the “keep alive” power is 144 hours (6 days to include margin). A 12v ArLSEP will use rechargeable LiFePO4 cells, which are common in the Electric Vehicle (EV) industry. With a current of 5 amps and a 125 watt system, the mass is about 90kg. The comm. system and structure adds another 10kg, thus the “Central Station” is approximately 100kg. The solar power is collected on four arrays (each 2m above the surface), and the entire ArLSEP is designed to stow in a 2m x 1m x 1m volume. The experiment and instrument design will vary for each installation and add mass to the total (although they are expected to fit within the 2m3 volume). Seismic wave generation will likely not be provided with mortars, thus an electric “thumper” will be required. Active instruments such as imaging systems and sensing instruments will benefit from the additional power and communication capability provided by ArLSEP. Passive systems such as retroreflectors, witness plates, and cosmic dust collectors can be added to either the landing vehicle and/or the ArLSEP. With repeated HLS missions to the same human site, the ArLSEP can be expanded and easily maintained for long duration science collection on the lunar surface.

ALSEP↗