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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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Planetary missions of the future will have increasingly greater energy requirements due to the desire for in-situ investigations and faster flight times. Solar electric propulsion provides a means of more effectively accomplishing these types of missions.
Development is underway on a unique high-voltage, high energy solar concentrator array called Stretched Lens Array Square-Rigger (SLASR) for direct drive electric propulsion. The SLASR performance attributes closely match the critical needs of solar electric propulsion (SEP) systems, which may be used for space tugs to fuel efficiently transport cargo from low earth orbit (LEO) to low lunar orbit (LLO), in support of NASA's robotic and human exploration missions. Later SEP systems may similarly transport cargo from the earth-moon neighborhood to the Mars neighborhood. This paper will describe the SLASR technology, discuss SLASR developments and ground testing, and outline plans for future SLASR technology maturation.
The thermal control requirements consist of functional requirements related to the various mission phase natural environments, operational requirements of induced power loadings by the solar electric propulsion stage subsystems, and design temperature limits for performance and reliability. The design approach utilizes passive thermal control techniques combining insulation, surface coatings, and sunshields with thermostatically controlled louvers. Heaters are used to regulate certain temperatures for extreme conditions. Details regarding the thruster array thermal control design are discussed, giving attention to the parameters used in the mathematical model, questions of conductive coupling, and thruster estimated power distribution.
Primitive bodies are exciting targets for exploration as they provide clues to the early Solar system conditions and dynamical evolution. The two moons of Mars are particularly interesting because of their proximity to an astrobiological target. However, after four decades of Mars exploration, their origin and nature remain enigmatic. In addition, when considering the long-term objectives of the flexible path for the potential human exploration to Mars, Phobos and Deimos present exciting intermediate opportunities without the complication and expense of landing and ascending from the surface. As interest in these targets for the next frontier of human exploration grows, characterization missions designed specifically to examine surface properties, landing environments, and surface mapping prior to human exploration are becoming increasingly important. A precursor mission concept of this sort has been developed using two identical spacecraft designed from low cost, flight proven and certified off-the-shelf component and utilizing Solar Electric Propulsion (SEP) to orbit both targets as secondary payloads launched aboard any NASA or GTO launch. This precursor mission has the potential to address both precursor measurements that are strategic knowledge gaps and decadal science, including soil physical properties at the global and local (human) scale and the search for in situ resources.
Human and robotic exploration beyond Low Earth Orbit (LEO) will require enabling capabilities that are efficient, affordable, and reliable. Solar Electric Propulsion (SEP) is highly advantageous because of its favorable in-space mass transfer efficiency compared to traditional chemical propulsion systems. The NASA studies have demonstrated that this advantage becomes highly significant as missions progress beyond Earth orbit. Recent studies of human exploration missions and architectures evaluated the capabilities needed to perform a variety of human exploration missions including missions to Near Earth Objects (NEOs). The studies demonstrated that SEP stages have potential to be the most cost effective solution to perform beyond LEO transfers of high mass cargoes for human missions. Recognizing that these missions require power levels more than 10X greater than current electric propulsion systems, NASA embarked upon a progressive pathway to identify critical technologies needed and a plan for an incremental demonstration mission. The NASA studies identified a 30kW class demonstration mission that can serve as a meaningful demonstration of the technologies, operational challenges, and provide the appropriate scaling and modularity required. This paper describes the planning options for a representative demonstration 30kW class SEP mission.
Development is underway on a unique high-voltage, high-energy solar concentrator array called Stretched Lens Array Square-Rigger (SLASR) for direct drive electric propulsion. The SLASR performance attributes closely match the critical needs of solar electric propulsion (SEP) systems, which may be used for space tugs to fuel-efficiently transport cargo from low earth orbit (LEO) to low lunar orbit (LLO), in support of NASA s robotic and human exploration missions. Later SEP systems may similarly transport cargo from the earth-moon neighborhood to the Mars neighborhood. This paper will describe the SLASR technology, discuss SLASR developments and ground testing, and outline plans for future SLASR technology maturation.
A slow flyby of the periodic comet, Encke, prior to its perihelion passage in 1980 has been shown to be a feasible mission for a solar electric propulsion spacecraft using a moderate power level of around 15 kw. The mission is accomplished in slightly less than two years with a planned miss distance and flyby speed of 1000 km and 4 km/s respectively. The results of a performance tradeoff study are presented together with a description of the nominal trajectory. The trajectories used in these mission studies were calculated using an updated model of the 30 cm mercury bombardment thruster's performance. The trajectory simulations include optimum staging of operating thrusters as a function of input power level constraints and limits on the variation of thruster specific impulse and efficiency as a function of thruster input power.
Three approaches are presented for packaging the elements of a 30 cm ion thruster subsystem into a modular thrust subsystem. The individual modules, when integrated into a conceptual solar electric propulsion module are applicable to a multimission set of interplanetary flights with the space shuttle interim upper stage as the launch vehicle. The emphasis is on the structural and thermal integration of the components into the modular thrust subsystems. Thermal control for the power processing units is either by direct radiation through louvers in combination with heat pipes or an all heat pipe system. The propellant storage and feed system and thruster gimbal system concepts are presented. The three approaches are compared on the basis of mass, cost, testing, interfaces, simplicity, reliability, and maintainability.
Three approaches are presented for packaging the elements of a 30 cm ion thrustor subsystem into a modular thrust subsystem. The individual modules, when integrated into a conceptual solar electric propulsion module are applicable to a multimission set of interplanetary flights with the Space Shuttle/Interim Upper Stage as the launch vehicle. The emphasis is on the structural and thermal integration of the components into the modular thrust subsystems. Thermal control for the power processing units is either by direct radiation through louvers in combination with heat pipes of an all heat pipe system. The propellant storage and feed system and thrustor gimbal system concepts are presented. The three approaches are compared on the basis of mass, cost, testing, interfaces, simplicity, reliability, and maintainability.
Highly propellant-efficient electric propulsion is being combined with advanced solar power technology to provide a non-nuclear transportation option for the human exploration of Mars. By virtue of its high specific impulse, electric propulsion offers a greater change in spacecraft velocity for each pound of propellant than do conventional chemical rockets. As a result, a mission to Mars based on solar electric propulsion (SEP) would require fewer heavy-lift launches than a traditional all-chemical space propulsion scenario would. Performance, as measured by mass to orbit and trip time, would be comparable to the NASA design reference mission for human Mars exploration, which utilizes nuclear thermal propulsion; but it would avoid the issues surrounding the use of nuclear reactors in space.
Beginning in 2014, the National Aeronautics and Space Administration (NASA) Space Technology Mission Directorate (STMD) began a project to increase the state of the art for the Hall-Effect Solar Electric Propulsion (SEP) technology. The Technology Demonstration Missions (TDM) Program Office provides programmatic oversight of SEP with the project being led by Glenn Research Center, supported by the Jet Propulsion Laboratory and development, qualification & flight hardware all provided by Aerojet Rocketdyne (AR). This technology was pursued as the propulsion system for the Asteroid Redirect and Robotic Mission (ARRM). While the concept was originally slated for ARRM, it was realigned to Gateway to support future Moon to Mars objectives. The Gateway lunar station was established and will play a key role in NASA’s Artemis Program which utilizes collaboration with the Canadian Space Agency (CSA), the European Space Agency (ESA) and the Japanese Space Agency (JAXA). The 12-kW hardware will be the primary propulsion for the Gateway element, Power & Propulsion Element (PPE), from Low Earth Orbit to a Near Rectilinear Halo Orbit around the Moon. Initial efforts began with utilizing Technology Development Units (TDU) built by NASA as the predecessor to the contract activity with AR. Over the past few years, AR has built Engineering Test Units (ETU), Engineering Development Units (EDU) and qualification & flight models (QM and FM, respectively). With requirement changes from the AARM mission to PPE, the joint team has modified the design and has started qualification activities for environment testing (shock, vibe) at multiple facilities in the US leading to eventual life testing. This paper will explore the various design changes, system modeling and the latest status for qualification testing.
Beginning in 2014, the National Aeronautics and Space Administration (NASA) Space Technology Mission Directorate (STMD) began a project to increase the state of the art for the Hall-Effect Solar Electric Propulsion (SEP) technology. The Technology Demonstration Missions (TDM) Program Office provides programmatic oversight of SEP with the project being led by Glenn Research Center, supported by the Jet Propulsion Laboratory and development, qualification & flight hardware all provided by Aerojet Rocketdyne (AR). This technology was pursued as the propulsion system for the Asteroid Redirect and Robotic Mission (ARRM). While the concept was originally slated for ARRM, it was realigned to Gateway to support future Moon to Mars objectives. The Gateway lunar station was established and will play a key role in NASA’s Artemis Program which utilizes collaboration with the Canadian Space Agency (CSA), the European Space Agency (ESA) and the Japanese Space Agency (JAXA). The 12-kW hardware will be the primary propulsion for the Gateway element, Power & Propulsion Element (PPE), from Low Earth Orbit to a Near Rectilinear Halo Orbit around the Moon. Initial efforts began with utilizing Technology Development Units (TDU) built by NASA as the predecessor to the contract activity with AR. Over the past few years, AR has built Engineering Test Units (ETU), Engineering Development Units (EDU) and qualification & flight models (QM and FM, respectively). With requirement changes from the AARM mission to PPE, the joint team has modified the design and has started qualification activities for environment testing (shock, vibe) at multiple facilities in the US leading to eventual life testing. This paper will explore the various design changes, system modeling and the latest status for qualification testing.
Developing a capability for the assembly of large space structures has the potential to increase the capabilities and performance of future space missions and spacecraft while reducing their cost. One such application is a megawatt-class solar electric propulsion (SEP) tug, representing a critical transportation ability for the NASA lunar, Mars, and solar system exploration missions. A series of robotic assembly experiments were recently completed at Langley Research Center (LaRC) that demonstrate most of the assembly steps for the SEP tug concept. The assembly experiments used a core set of robotic capabilities: long-reach manipulation and dexterous manipulation. This paper describes cross-cutting capabilities and technologies for in-space assembly (ISA), applies the ISA approach to a SEP tug, describes the design and development of two assembly demonstration concepts, and summarizes results of two sets of assembly experiments that validate the SEP tug assembly steps.
Program SEPSIM is a FORTRAN program which performs deployment, servicing, and retrieval missions to synchronous equatorial orbit using a space tug with a continuous low thrust upper stage known as a solar electric propulsion stage (SEPS). The SEPS ferries payloads back and forth between an intermediate orbit and synchronous orbit, and performs the necessary servicing maneuvers in synchronous orbit. The tug carries payloads between the orbiter and the intermediate orbit, deploys fully fueled SEPS vehicles, and retrieves exhausted SEPS vehicles when, and if, required. The program is presently contained in subroutine form in the Logistical On-orbit VEhicle Servicing (LOVES) Program, but can also be run independently with the addition of a simple driver program.
Summary of a feasibility study of CARD mission/system concepts for comet Encke and asteroid Eros missions in the late 1970s. A common planetary vehicle employing a modular SEP (solar electric propulsion) system and a direct rendezvous/docking mode with a staged science/sampling module appears feasible based on 1973 technology. The SEP system utilizes 3.5-kW, 3500-sec Isp mercury ion thrusters and rollout solar arrays sized at 36 (Encke) and 13 (Eros) kW based on the Titan III family launch vehicles. Science payloads, sampling concepts, and supporting spacecraft subsystems are defined. Problem areas are identified, and programmatic considerations are discussed.
This paper summarizes a portion of the planetary mission analysis results of past and present studies conducted by Rockwell International for NASA-MSFC (Contract NAS8-27360) dealing with the feasibility of a Solar Electric Propulsion Stage (SEPS). The SEPS is envisioned as an upper stage of a transportation system capable of delivering either separable payload spacecraft or attached science packages to various planetary targets. The purpose of the paper is to demonstrate that, from a payload performance capability standpoint, a common SEP Stage can deliver various payloads to a host of planetary targets including inner and outer planets, asteroids, and comets.
This paper documents the results of an investigation of human Mars mission architectures that leverage near-term technology investments and infrastructures resulting from the planned Asteroid Redirect Robotic Mission (ARRM), including high-power Solar Electric Propulsion (SEP) and a human presence in Lunar Distant Retrograde Orbit (LDRO). The architectures investigated use a combination of SEP and chemical propulsion elements. Through this combination of propulsion technologies, these architectures take advantage of the high efficiency SEP propulsion system to deliver cargo, while maintaining the faster trip times afforded by chemical propulsion for crew transport. Evolved configurations of the Asteroid Redirect Vehicle (ARV) are considered for cargo delivery. Sensitivities to SEP system design parameters, including power level and propellant quantity, are presented. For the crew delivery, liquid oxygen and methane stages were designed using engines common to future human Mars landers. Impacts of various Earth departure orbits, Mars loiter orbits, and Earth return strategies are presented. The use of the Space Launch System for delivery of the various architecture elements was also investigated and launch vehicle manifesting, launch scheduling and mission timelines are also discussed. The study results show that viable Mars architecture can be constructed using LDRO and SEP in order to take advantage of investments made in the ARRM mission.
This paper presents study results quantifying the benefits of higher voltage, electric power system designs for a typical solar electric propulsion spacecraft Earth orbiting mission. A conceptual power system architecture was defined and design points were generated for system voltages of 28-V, 50-V, 120-V, and 300-V using state-of-the-art or advanced technologies. A 300-V 'direct-drive' architecture was also analyzed to assess the benefits of directly powering the electric thruster from the photovoltaic array without up-conversion. Fortran and spreadsheet computational models were exercised to predict the performance and size power system components to meet spacecraft mission requirements. Pertinent space environments, such as electron and proton radiation, were calculated along the spiral trajectory. In addition, a simplified electron current collection model was developed to estimate photovoltaic array losses for the orbital plasma environment and that created by the thruster plume. The secondary benefits of power system mass savings for spacecraft propulsion and attitude control systems were also quantified. Results indicate that considerable spacecraft wet mass savings were achieved by the 300-V and 300-V direct-drive architectures.