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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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At least 487 records · Page 27
All sourcesNASA NTRSDOE OSTI
NASA NTRS2023

Combined 1-MW Solar Electric and Chemical Propulsion for Crewed Mars Missions

The current Moon to Mars Architecture defines functional capabilities that are essential to achieving NASA’s specific exploration goals and objectives. The transportation system goals and objectives or functional capabilities do not prescribe a solution, leaving an open tradespace to be explored. NASA is working to identify power and propulsion technologies that enable feasible transportation options for crewed missions to Mars. During NASA’s Strategic Analysis Cycle 2021 (SAC21), nuclear enabled spacecraft concepts were analyzed for moderate duration 850-day Earth-Mars roundtrip mission2. During SAC22, the transportation tradespace was expanded to understand the feasibility of using different propulsion systems for crewed Mars missions using a moderate duration reference mission as a comparison point. This paper explores the mission analysis performed for a conceptual MW-class hybrid Solar Electric Propulsion (SEP)-chemical propulsion system (SEP-Chem) spacecraft optimized for an 850-day crewed Mars mission for the 2039 opportunity.

SEP-Chem↗
NASA NTRS2008

NASA's Evolutionary Xenon Thruster: The NEXT Ion Propulsion System for Solar System Exploration

This viewgraph presentation reviews NASA s Evolutionary Xenon Thruster (NEXT) Ion Propulsion system. The NEXT project is developing a solar electric ion propulsion system. The NEXT project is advancing the capability of ion propulsion to meet NASA robotic science mission needs. The NEXT system is planned to significantly improve performance over the state of the art electric propulsion systems, such as NASA Solar Electric Propulsion Technology Application Readiness (NSTAR). The status of NEXT development is reviewed, including information on the NEXT Thruster, the power processing unit, the propellant management system (PMS), the digital control interface unit, and the gimbal. Block diagrams NEXT system are presented. Also a review of the lessons learned from the Dawn and NSTAR systems is provided. In summary the NEXT project activities through 2007 have brought next-generation ion propulsion technology to a sufficient maturity level.

Pencil, Eric J.↗
NASA NTRS2011

Strategic Implications of Human Exploration of Near-Earth Asteroids

The current United States Space Policy [1] as articulated by the White House and later confirmed by the Congress [2] calls for [t]he extension of the human presence from low-Earth orbit to other regions of space beyond low-Earth orbit will enable missions to the surface of the Moon and missions to deep space destinations such as near-Earth asteroids and Mars. Human exploration of the Moon and Mars has been the focus of numerous exhaustive studies and planning, but missions to Near-Earth Asteroids (NEAs) has, by comparison, garnered relatively little attention in terms of mission and systems planning. This paper examines the strategic implications of human exploration of NEAs and how they can fit into the overall exploration strategy. This paper specifically addresses how accessible NEAs are in terms of mission duration, technologies required, and overall architecture construct. Example mission architectures utilizing different propulsion technologies such as chemical, nuclear thermal, and solar electric propulsion were formulated to determine resulting figures of merit including number of NEAs accessible, time of flight, mission mass, number of departure windows, and length of the launch windows. These data, in conjunction with what we currently know about these potential exploration targets (or need to know in the future), provide key insights necessary for future mission and strategic planning.

Drake, Bret G.↗
NASA NTRS2015

SEP Mission Design Space for Mars Orbiters

The advancement of solar-electric propulsion (SEP) technologies and larger, light-weight solar arrays offer a tremendous advantage to Mars orbiters in terms of both mass and timeline flexibility. These advantages are multiplied for round-trip orbiters (e.g. potential Mars sample return) where a large total Delta V would be required. In this paper we investigate the mission design characteristics of mission concepts utilizing various combinations and types of SEP thrusters, solar arrays, launch vehicles, launch dates, arrival dates, etc. SEP allows for greater than 50% more mass delivered and launch windows of months to years. We also present the SEP analog to the ballistic Porkchop plot - the "Bacon" plot.

Woolley, Ryan C.↗
NASA NTRS1973

Electric propulsion status and development plans - NASA programs

The current status of the NASA auxiliary and primary electric propulsion programs is reviewed. Plans for continuation of these programs and future target goals are also outlined. In auxiliary propulsion, progress on the ion engine flight experiment for the Application Technology Satellite (ATS)-F launch in 1974 to demonstrate North-South station keeping capability is reviewed. In addition, work on small pulsed plasma thrusters and hydrazine resistojets for attitude control applications are also described. In primary propulsion, the effort aimed at demonstrating the technology for solar electric propulsion is presented.

Lazar, J.↗
NASA NTRS2015

Comparison of Optimal Small Spacecraft Micro Electric Propulsion Technologies for Mission Opportunities

The goal of this paper is to explore the mission opportunities that are uniquely enabled by U-class Solar Electric Propulsion (SEP) technologies. Small SEP thrusters offers significant advantages relative to existing technologies and will revolutionize the class of mission architectures that small spacecraft can accomplish by enabling trajectory maneuvers with significant change in velocity requirements and reaction wheel-free attitude control. This paper aims to develop and apply a common system-level modeling framework to evaluate these thrusters for relevant upcoming mission scenarios, taking into account the mass, power, volume, and operational constraints of small highly-constrained missions. We will identify the optimal technology for broad classes of mission applications for different U-class spacecraft sizes and provide insights into what constrains the system performance to identify technology areas where improvements are needed.

Spangelo, Sara↗
NASA NTRS2017

13kW Advanced Electric Propulsion Flight System Development and Qualification

The next phase of robotic and human deep space exploration missions is enhanced by high performance, high power solar electric propulsion systems for large-scale science missions and cargo transportation. Aerojet Rocketdynes Advanced Electric Propulsion System (AEPS) program is completing development, qualification and delivery of five flight 13.3kW EP systems to NASA. The flight AEPS includes a magnetically-shielded, long-life Hall thruster, power processing unit (PPU), xenon flow controller (XFC), and intrasystem harnesses. The Hall thruster, originally developed and demonstrated by NASAs Glenn Research Center and the Jet Propulsion Laboratory, operates at input powers up to 12.5kW while providing a specific impulse over 2600s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95 to 140V, consistent with operation beyond the orbit of Mars. The integrated system is continuously throttleable between 3 and 13.3kW. The program has completed the system requirement review; the system, thruster, PPU and XFC preliminary design reviews; development of engineering models, and initial system integration testing. This paper will present the high power AEPS capabilities, overall program and design status and the latest test results for the 13.3kW flight system development and qualification program.

Jackson, Jerry↗
NASA NTRS1982

Electric propulsion research and technology in the United States

Near-, mid-, and long-term technology goals for space electric propulsion systems are reviewed. Technological readiness has been demonstrated for 8 cm, 5 mN, and 30 cm, 30 mN electrostatic thrusters, with major use seen for GEO communications satellites in the near-term, and space station orbit adjustments later. Ion thrusters and/or MPD thrusters are projected to become viable if a space nuclear reactor system is operational in the 1990s, allowing the transport of thousands of kilograms to the outer planets. Basic research is proceeding on the electrothermal propulsion concept to provide resistojet thrusters suitable for a space station by 1986, although the program is hindered by insufficient funding. A flight test for the ion auxiliary propulsion system is detailed, and test results of the solar electric rocket (SERT II) are reported. Particular note is made of the progress on the solar electric propulsion system for the Shuttle and the Nuclear electric propulsion system for thermal-to-electric conversion for mid-1990s applications.

Hudson, W. R.↗
NASA NTRSDate not supplied

Power and Propulsion Element for Gateway

NASA is taking on the bold objective of returning humans to the Moon by 2024. As part of this effort, the Gateway is being constructed to serve as a command center and way-point in cis-lunar space for lunar exploration missions. The Power and Propulsion Element (PPE) is the first element of Gateway to be launched and will generate all the power and propulsion required over its 15-year service life. To achieve these goals, high-power solar arrays and solar electric propulsion will be implemented and demonstrated. The electrical power system (EPS) of Gateway, prominently featuring lightweight Roll Out Solar Arrays (ROSA) with state-of-the-art Z4J solar cells and a central hinge, will be sized for an end-of-life (EOL) power management and distribution capability of over 55 kilowatts (kW). The solar array circuits are being designed in a modular manner with scalability and extensibility to a 300 kW-class mission. NASA Glenn Research Center is managing the PPE contract on which Maxar Technologies, Inc. is responsible for fabricating and demonstrating the spacecraft. This manuscript provides a summary of the electrical power system, with additional discussion of the ROSA power generation.

Gateway↗
NASA NTRS2015

Integrated Vehicle and Trajectory Design of Small Spacecraft with Electric Propulsion for Earth and Interplanetary Missions

This paper investigates the feasibility of Earth-transfer and interplanetary mission architectures for miniaturized spacecraft using emerging small solar electric propulsion technologies. Emerging small SEP thrusters offer significant advantages relative to existing technologies and will enable U-class systems to perform trajectory maneuvers with significant Delta V requirements. The approach in this paper is unique because it integrates trajectory design with vehicle sizing and accounts for the system and operational constraints of small U-class missions. The modeling framework includes integrated propulsion, orbit, energy, and external environment dynamics and systems-level power, energy, mass, and volume constraints. The trajectory simulation environment models orbit boosts in Earth orbit and flyby and capture trajectories to interplanetary destinations. A family of small spacecraft mission architectures are studied, including altitude and inclination transfers in Earth orbit and trajectories that escape Earth orbit and travel to interplanetary destinations such as Mercury, Venus, and Mars. Results are presented visually to show the trade-offs between competing performance objectives such as maximizing available mass and volume for payloads and minimizing transfer time. The results demonstrate the feasibility of using small spacecraft to perform significant Earth and interplanetary orbit transfers in less than one year with reasonable U-class mass, power, volume, and mission durations.

Spangelo, Sara↗
NASA NTRS1995

Advanced Propulsion for Geostationary Orbit Insertion and North-South Station Keeping

Solar electric propulsion (SEP) technology is currently being used for geostationary satellite station keeping to increase payload mass. Analyses show that advanced electric propulsion technologies can be used to obtain additional increases in payload mass by using these same technologies to perform part of the orbit transfer. In this work three electric propulsion technologies are examined at two power levels for an Atlas 2AS class spacecraft. The on-board chemical propulsion apogee engine fuel is reduced to allow the use of electric propulsion. A numerical optimizer is used to determine the chemical burns which will minimize the electric propulsion transfer time. Results show that for a 1550 kg Atlas 2AS class payload, increases in net mass (geostationary satellite mass less wet propulsion system mass) of 150 to 800 kg are possible using electric propulsion for station keeping, advanced chemical engines for part of the transfer, and electric propulsion for the remainder of the transfer. Trip times are between one and four months.

Oleson, Steven R.↗
NASA NTRS1997

Advanced Propulsion for Geostationary Orbit Insertion and North-South Station Keeping

Solar electric propulsion technology is currently being used for geostationary satellite station keeping. Analyses show that electric propulsion technologies can be used to obtain additional increases in payload mass by using them to perform part of the orbit transfer. Three electric propulsion technologies are examined at two power levels for geostationary insertion of an Atlas IIAS class spacecraft. The onboard chemical propulsion apogee engine fuel is reduced in this analysis to allow the use of electric propulsion. A numerical optimizer is used to determine the chemical burns that will minimize the electric propulsion transfer times. For a 1550-kg Atlas IIAS class payload, increases in net mass (geostationary satellite mass less wet propulsion system mass) of 150-800 kg are enabled by using electric propulsion for station keeping, advanced chemical engines for part of the transfer, and electric propulsion for the remainder of the transfer. Trip times are between one and four months.

Oleson, Steven R.↗
NASA NTRS1973

Review of Kaufman thruster development at the Lewis Research Center - 1973

Work on Kaufman thruster development completed during the years 1971 and 1972 is reviewed. Thrusters tested have ranged in size from 2.5-cm to 150-cm diameters, in thrust from 0.4 to 4300 mN, and in power from 0.03 to 203 kW. A 2.5-cm thruster was briefly tested and found to have surprisingly high thruster efficiency. Emphasis is placed on thruster system reliability and lifetime as previous work has increased thruster efficiency to a high level. Work also proceeds on definition of thruster-spacecraft interactions. Major R&D efforts are directed at present into two areas of thruster size: a 5-cm to 8-cm diameter thruster to be used for station keeping and attitude control of geosynchronous spacecraft; and a 30-cm diameter thruster to be used for primary propulsion in a 3- to 7-thruster array for solar electric propulsion of interplanetary spacecraft.

Kerslake, W. R.↗
NASA NTRS1976

Development of a 30-cm ion thruster thermal-vacuum power processor

The 30-cm Hg electron-bombardment ion thruster presently under development has reached engineering model status and is generally accepted as the prime propulsion thruster module to be used on the earliest solar electric propulsion missions. This paper presents the results of a related program to develop a transistorized 3-kW Thermal-Vacuum Breadboard (TVBB) Power Processor for this thruster. Emphasized in the paper are the implemented electrical and mechanical designs as well as the resultant system performance achieved over a range of test conditions. In addition, design modifications affording improved performance are identified and discussed.

Herron, B. G.↗
NASA NTRS1995

Development Status of the NSTAR Ion Propulsion System Power Processor

A 0.5-2.3 kW xenon ion propulsion system is presently being developed under the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) program. This propulsion system includes a 30 cm diameter xenon ion thruster, a Digital Control Interface Unit, a xenon feed system, and a power processing unit (PPU). The PPU consists of the power supply assemblies which operate the thruster neutralizer, main discharge chamber, and ion optics. Also included are recycle logic and a digital microcontroller. The neutralizer and discharge power supplies employ a dual use configuration which combines the functions of two power supplies into one, significantly simplifying the PPU. Further simplification was realized by implementing a single thruster control loop which regulates the beam current via the discharge current. Continuous throttling is possible over a 0.5-2.3 kW output power range. All three power supplies have been fabricated and tested with resistive loads, and have been combined into a single breadboard unit with the recycle logic and microcontroller. All line and load regulation test results show the power supplies to be within the NSTAR flight PPU specified power output of 1.98 kW. The overall efficiency of the PPU, calculated as the combined efficiencies of the power supplies and controller, at 2.3 kW delivered to resistive loads was 0.90. The component was 6.16 kg. Integration testing of the neutralizer and discharge power supplies with a functional model thruster revealed no issues with discharge ignition or steady state operation.

Hamley, John A.↗
NASA NTRS1975

SEPS solar array design and technology evaluation

The technology developments required and a preliminary design of a lightweight 25 kW solar array for the solar electric propulsion stage (SEPS) have been defined. The requirements for a 65 W/Kg SEPS solar array system requires significant component weight reductions over present state-of-the-art flexible solar arrays in both electrical and structural-mechanical designs. A requirement for operation from 0.3 au to 6.0 au presents a wide range of temperature environments as well as severe combined thermal/vacuum/UV radiation environments. Additional requirements are capability for partial array retraction operation, and capability for full retraction and automatic preloading for survival of the Shuttle reentry environment. An assessment of current lightweight flexible solar array technology is made against the SEPS solar array requirements and new technology requirements are defined. A preliminary design and the operating characteristics of a flat-fold solar array system meeting the SEPS requirements is presented. A full-width, 10-ft-tall functional array model, including representative welded electrical modules and a model astromast, was fabricated and tested.

Elms, R. V., Jr.↗
NASA NTRS2004

SEP Mission to Titan NEXT Aerocapture In-Space Propulsion (Quicktime Movie)

The ion thruster is one of the most promising solar electric propulsion (SEP) technologies to support future Outer Planet missions (place provided link below here) for NASA's Office of Space Science. Typically, ion thrusters are used in high Isp- low thrust applications that require long lifetimes, as well as, higher efficiency over state-of-the-art chemical propulsion systems.Today, the standard for ion thrusters is the SEP Technology Application Readiness (NSTAR) thruster. Jet Propulsion Laboratory's (JPL's) extended life test (ELT) of the DS 1 flight spare NSTAR thruster began in October 1998. This test successfully demonstrated lifetime of the NSTAR flight spare thruster, which will provide a solid basis for selection of ion thrusters for future Code S missions. The NSTAR ELT was concluded on June 30,2003 after 30,352 hours. The purpose of the Next Generation Ion (NGI) activities is to advance Ion propulsion system technologies through the development of NASA's Evolutionary Xenon Thruster (NEXT). The goal of NEXT is to more than double the power capability and lifetime throughput (the total amount of propellant which can be processed) while increasing the Isp by 30% and the thrust by 120%.

Baggett, Randy↗
NASA NTRS2011

Power and Propulsion System Design for Near-Earth Object Robotic Exploration

Near-Earth Objects (NEOs) are exciting targets for exploration; they are relatively easy to reach but relatively little is known about them. With solar electric propulsion, a vast number of interesting NEOs can be reached within a few years and with extensive flexibility in launch date. An additional advantage of electric propulsion for these missions is that a spacecraft can be small, enabling a fleet of explorers launched on a single vehicle or as secondary payloads. Commercial, flight-proven Hall thruster systems have great appeal based on their performance and low cost risk, but one issue with these systems is that the power processing units (PPUs) are designed for regulated spacecraft power architectures which are not attractive for small NEO missions. In this study we consider the integrated design of power and propulsion systems that utilize the capabilities of existing PPUs in an unregulated power architecture. Models for solar array and engine performance are combined with low-thrust trajectory analyses to bound spacecraft design parameters for a large class of NEO missions, then detailed array performance models are used to examine the array output voltage and current over a bounded mission set. Operational relationships between the power and electric propulsion systems are discussed, and it is shown that both the SPT-100 and BPT-4000 PPUs can perform missions over a solar range of 0.7 AU to 1.5 AU - encompassing NEOs, Venus, and Mars - within their operable input voltage ranges. A number of design trades to control the array voltage are available, including cell string layout, array offpointing during mission operations, and power draw by the Hall thruster system.

Near Earth Object (NEO)↗