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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 343 records · Page 19
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NASA NTRS2010

TRL Assessment of Solar Sail Technology Development Following the 20-Meter System Ground Demonstrator Hardware Testing

The NASA In-Space Propulsion Technology (ISPT) Projects Office sponsored two separate, independent solar sail system design and development demonstration activities during 2002-2005. ATK Space Systems of Goleta, CA was the prime contractor for one development team and L' Garde, Inc. of Tustin, CA was the prime contractor for the other development team. The goal of these activities was to advance the technology readiness level (TRL) of solar sail propulsion from 3 towards 6 by the year 2006. Component and subsystem fabrication and testing were completed successfully, including the ground deployment of 10-meter and 20-meter demonstration hardware systems under vacuum conditions. The deployment and structural testing of the 20-meter solar sail systems was conducted in the 30 meter diameter Space Power Facility thermal-vacuum chamber at NASA Glenn Plum Brook in April though August, 2005. This paper will present the results of the TRL assessment following the solar sail technology development activities associated with the design, development, analysis and testing of the 20-meter system ground demonstrators.

Young, Roy M.↗
NASA NTRS2014

Overview of the Development of the Solar Electric Propulsion Technology Demonstration Mission 12.5-kW Hall Thruster

NASA is developing mission concepts for a solar electric propulsion technology demonstration mission. A number of mission concepts are being evaluated including ambitious missions to near Earth objects. The demonstration of a high-power solar electric propulsion capability is one of the objectives of the candidate missions under consideration. In support of NASA's exploration goals, a number of projects are developing extensible technologies to support NASA's near and long term mission needs. Specifically, the Space Technology Mission Directorate Solar Electric Propulsion Technology Demonstration Mission project is funding the development of a 12.5-kilowatt magnetically shielded Hall thruster system to support future NASA missions. This paper presents the design attributes of the thruster that was collaboratively developed by the NASA Glenn Research Center and the Jet Propulsion Laboratory. The paper provides an overview of the magnetic, plasma, thermal, and structural modeling activities that were carried out in support of the thruster design. The paper also summarizes the results of the functional tests that have been carried out to date. The planned thruster performance, plasma diagnostics (internal and in the plume), thermal, wear, and mechanical tests are outlined.

High power long life hall thrusters↗
NASA NTRS1975

Determination of the extent of ion thruster efflux deposition on spacecraft surfaces from the SERT II flight thermal data

The flight thermal data from the SERT II satellite, the only electric propulsion mission with an extensive thruster operational history, is reviewed specifically to see if there is any evidence of contamination that could be attributed to the 5860 hours of mercury bombardment ion thruster operation. This evaluation shows that the only evidence of deposition occurred on the contamination experiment solar cells, which are located at the edge of the thruster exhaust beam. There is no evidence of any deposition of ion thruster efflux on any other surface of the vehicle.

Stevens, N. J.↗
NASA NTRS1977

Thermal design of the IUE hydrazine auxiliary propulsion system

The International Ultraviolet Explorer is a large astronomical observatory scheduled to be placed in a three-axis stabilized synchronous orbit in the fourth quarter of 1977. The Hydrazine Auxiliary Propulsion System (HAPS) must perform a number of spacecraft maneuvers to achieve a successful mission. This paper describes the thermal design which accomplishes temperature control between 5 and 65 C for all orbital conditions by utilizing multilayer insulation and commandable component heaters. A primary design criteria was the minimization of spacecraft power by the selective use of the solar environment. The thermal design was carefully assessed and verified in both spacecraft thermal balance and subsystem solar simulation testing.

Joseph T. Skladany↗
NASA NTRSDate not supplied

BioSentinel: Leading the Way for Deep Space CubeSat Missions

Flagship science missions are not alone in Deep Space thanks to BioSentinel, a 6U spacecraft launched on Artemis-1. BioSentinel is one of the longest operating CubeSats beyond cislunar space. The subsystems and COTS components of the BioSentinel bus are a template for future deep space missions, and the lessons learned from over a year of operations will enable improved performance for the next missions. BioSentinel achieved its unprecedented performance for an SLS secondary payload due to preparation, planning, and a robust design. Pre-launch antenna and interface testing with both DSN and ESA confirmed command and data pathways and allowed for operational flexibility in the critical early hours post-deployment. Mission Operations simulations prior to launch identified potential risks and primed operators to respond in flight, preparing the team to react quickly to successfully detumble the spacecraft and enter a power-positive state. The spacecraft would not have survived without the inclusion of the trailblazing 3D-printed composite cold gas propulsion system. The non-standard tank geometry enabled efficient use of the limited space available in the CubeSat, as well as the capability to detumble the spacecraft and manage momentum, while providing sufficient margin to execute potential delta-V maneuvers. The Iris radio has operated for over 18 months with no significant issues. Initial Iris performance estimates have been accurate throughout the mission. BioSentinel continues to collect data on thermal conditions and to validate our performance models with real-world knowledge. We have received exemplary support from our DSN partners. Following the conclusion of the primary science mission, the Linear Energy Transfer (LET) Spectrometer continued to collect solar and galactic radiation data from its location in heliocentric orbit. The free space dataset offered by the BioSentinel LET is a valuable source of data for model validation and future mission planning. As the spacecraft travels farther from Earth it is poised to provide longitudinally distributed measurements of solar particle events during solar maximum. The lessons learned from BioSentinel suggest key areas to enhance performance. The ability to upload modified flight software can increase the stability of memory management. Additional heaters in the propulsion system design have already proven successful on the Starling mission. Streamlining mission operations can reduce costs, increase data return, and better utilize DSN time. Enhancements such as these will facilitate reliable, long-duration deep space exploration using the proven BioSentinel 6U CubeSat bus.

BioSentinel↗
NASA NTRSDate not supplied

Early Progress Toward the Feasibility of the Centrifugal Nuclear Thermal Rocket

The Centrifugal Nuclear Thermal Rocket (CNTR) is a Nuclear Thermal Propulsion (NTP) concept designed to heat propellant directly by the reactor fuel. The primary difference between the CNTR concept and traditional NTP systems is that rather than using traditional solid fuel elements, the CNTR uses liquid fuel with the liquid contained in rotating cylinders by centrifugal force. If the concept can be successfully realized, the CNTR would have a high specific impulse (~1800 seconds) at high thrust, which may enable (i) viable near-term human Mars exploration by reducing round-trip times to 420 days and (ii) direct injection orbits for scientific missions to the Solar System outer planets and potentially Kuiper Belt objects. The CNTR could also use storable propellants such as ammonia, methane, or hydrazine at an Isp of ~900 seconds, enabling long-term in-space storage of a dormant system. Significant engineering challenges must be addressed to establish the technical viability of the CNTR. Research is presently underway to determine resolutions for these engineering challenges. In particular, research has begun on the analytical modeling and simulation of the two-phase heat transfer between the liquid metallic uranium fuel and the gaseous propellant. A paper was presented at the 2021 IAC which described these challenges and the study plan to address them. This paper describes the analytical and experimental progress to date toward resolving these challenges and establishing the engineering feasibility of the CNTR technology

High Performance Nuclear Thermal Propulsion NTP SN↗
NASA NTRS1991

Coproduction of volatiles and metals from extraterrestrial materials

Two main efforts in support of the general goals of SERC/culpr are presented. Investigations of processes for the coproduction of metals from extra-terrestrial materials in conjunction with plausible schemes for oxygen extraction continue. The principal emphasis was on the extraction and purification of iron from the ilmenite reduction process for oxygen, from the cathode metal deposits made in the magma electrolysis process for oxygen, and from native ferrous metal alloys on the moon and asteroids. All work on the separation and purification of ferrous metals was focussed upon the gaseous carbonyl process, a scheme that involves only temperatures attainable by passive thermal control. The exploration of a variety of schemes was initiated, involving the use of several different propulsion options and both propulsive and aerobraking capture at earth, for return of extraterrestrial resources to earth orbits. In addition, the search for new opportunities in space resource utilization continues. Examples include the continuation of work underway on: (1) the feasibility of locating solar power satellites in highly eccentric earth orbit; (2) the energetics of extracting the potential clean fusion fuel He-3 from the atmosphere for return to earth; and (3) the utility of a nuclear steam rocket (using non-terrestrial water as the working fluid) for transportation in the inner solar system.

Lewis, John S.↗
NASA NTRS1972

Solar electric propulsion breadboard thrust subsystem test results.

Work providing a firm basis for building an engineering model thrust subsystem (TSS) is described. A 1500-hr functional demonstration test revealed the impact of each element on TSS integration, operation, and future development. The primary conclusions are: accurate and specific definition of interfaces is essential to a timely system integration program. In addition to normal operating requirements, this definition must include requirements for preflight checkout, self-protection of all interfaces, capability expansion, failure modes, maximum input power, interelement power losses, noise and transient sensitivity and production, packaging for vacuum, thermal control, electrical grounding, cabling, mounting, and handling. In any future design, each phase of the TSS operation (startup, throttling, switching, etc.) should be considered in detail to anticipate and account for possible in-flight variations in element characteristics and performance.

Masek, T. D.↗
NASA NTRS1990

Project WISH: The Emerald City

When Project WISH (Wandering Interplanetary Space Harbor) was initiated as a multi-year project, several design requirements were specified. The space station must have a lifetime of at least 50 years, be autonomous and independent of Earth resources, be capable of traveling throughout the solar system within a maximum flight time of three years, and have a population of 500 to 1000 people. The purpose of the station is to provide a permanent home for space colonists and to serve as a service station for space missions. The orbital mechanics, propulsion system, vehicle dynamics and control, life support system, communication system, power system, and thermal system are discussed.

Source record↗
NASA NTRS1978

Extended performance solar electric propulsion thrust system design

A thrust system design has been established for an extended performance technology, 6.4 kW, 4800 sec specific impulse ion thruster. The configuration is comprised of ten thrusters configured with a power management and control subsystem in a modular thrust system design. The system design approach is an adaptation of that previously established for the baseline technology 2.7 kW, 3000 sec specific impulse ion thruster. The power management and control subsystem design includes a combination of individual electronics for each thruster and a set of electronics with redundancy that are common to all thrusters. The thermal dissipation from all electronics is removed with a common heat pipe/radiator assembly.

Cake, J. E.↗
NASA NTRS1984

Salt-gradient Solar Ponds: Summary of US Department of Energy Sponsored Research

The solar pond research program conducted by the United States Department of Energy was discontinued after 1983. This document summarizes the results of the program, reviews the state of the art, and identifies the remaining outstanding issues. Solar ponds is a generic term but, in the context of this report, the term "solar pond" refers specifically to salt-gradient solar pond. Several small research solar ponds have been built and successfully tested. Procedures for filling the pond, maintaining the gradient, adjusting the zone boundaries, and extracting heat have been developed. Theories and models have also been developed and verified. The major remaining unknowns or issues involve the physical behavior of large ponds; i.e., wind mixing of the surface, lateral range or reach of horizontally injected fluids, ground thermal losses, and gradient zone boundary erosion caused by pumping fluid for heat extraction. These issues cannot be scaled and must be studied in a large outdoor solar pond. This report has been subdivided into three parts. Part One presents the results of the DOE research program. The major contributing institutions include Solar Energy Research Institute, Los Alamos National Laboratory, and Jet Propulsion Laboratory. Part Two presents related independent research conducted by universities. Part Three is a selected bibliography with abstracts.

R.L. French↗
NASA NTRS2003

The Neptune/Triton Explorer Mission: A Concept Feasibility Study

Technological advances over the next 10 to 15 years promise to enable a number of smaller, more capable science missions to the outer planets. With the inception of miniaturized spacecraft for a wide range of applications, both in large clusters around Earth, and for deep space missions, NASA is currently in the process of redefining the way science is being gathered. Technologies such as 3-Dimensional Multi-Chip Modules, Micro-machined Electromechanical Devices, Multi Functional Structures, miniaturized transponders, miniaturized propulsion systems, variable emissivity thermal coatings, and artificial intelligence systems are currently in research and development, and are scheduled to fly (or have flown) in a number of missions. This study will leverage on these and other technologies in the design of a lightweight Neptune orbiter unlike any other that has been proposed to date. The Neptune/Triton Explorer (NExTEP) spacecraft uses solar electric earth gravity assist and aero capture maneuvers to achieve its intended target orbit. Either a Taurus or Delta-class launch vehicle may be used to accomplish the mission.

Esper, Jaime↗
NASA NTRS1975

Determination of the extent of ion thruster efflux

In the studies of proposed electric propulsion missions one of the areas of concern is the possible contamination of spacecraft instruments and thermal control surfaces by exhaust particles from an ion thruster. Vacuum tank tests were conducted in ground facilities to determine the extent of this deposition by thruster exhaust particles, but the application of these results to long term space missions is questionable. The flight thermal data from the SERT II satellite, the only electric propulsion mission with an extensive thruster operational history, was reviewed specifically to see if there is any evidence of contamination that could be attributed to the 5860 hours of mercury bombardment ion thruster operation. This evaluation of the flight data shows that the only evidence of deposition occurred on the contamination experiment solar cells which are located at the edge of the thruster exhaust beam. There is no evidence of any deposition of ion thruster efflux on any other surface of the satellite.

Stevens, N. J.↗
NASA NTRS2006

Material Interactions with Solar Wind Ion Environments

Solar wind composition is dominated by hydrogen (approx.96%) and helium (approx.3 to 4%) with a minor fraction (less than or equal to 1%) of heavy ions. Hydrogen (helium) ions impact spacecraft surfaces with energies from 0.5 to 5 keV (1.8 to 21 keV) due to variations in solar wind velocity from 300 km/s to 1000 km/sec with extremes of a few 10 s keV during periods of extremely high solar wind velocity exceeding 1000 km/sec. Mean impact energies are typically on the order of approximately 1 keV and 4 keV for hydrogen ions and helium ions, respectively. These energies are typically of the peak of the energy dependent light ion sputter yields for hydrogen and helium on many metals. In addition, light ions with kilovolt energies have been shown to produce blister (or exfoliation) damage to metal surfaces due to formation of high pressure gas bubbles within the materials when exposed to ion fluences on the order of 10(exp 16 to (10(exp 17 ions/sq cm. A number of spacecraft designs for current and future missions include gossamer polymer structures with thin metallic reflection coatings to shield instruments from the Sun or solar sail propulsion systems for use in a variety of locations in the inner solar system from 0.5 to 1 AU. In addition, there is interest in designing spacecraft for solar physics missions requiring operations as close to the Sun as 0.16 to 0.2 AU. Integrity of the metallic coatings is critical in many of these applications since degradation will result in modification of material thermal properties or exposure of polymers to solar UV photons which can compromise mission requirements. This paper will evaluate the relative contributions of sputtering and blister formation to material degradation in solar wind environments over a range of radial distances from the Sun to demonstrate where solar wind environments become important for materials selection. We will first review the physics and results from laboratory measurements of light ion sputtering, blistering, and exfoliation of metallic surfaces to establish the order of magnitude ion fluence required for significant surface damage. Solar wind ion fluence environments will then be evaluated due to variations in solar wind conditions as a function of solar cycle for varying distances from the Sun using models for radial variations in solar wind ion number density, temperature, and velocity to determine where sputtering and blistering is most likely to be an issue. Finally, ion fluence statistics for varying radial distances from the Sun will be shown to establish the mission duration and radial distances from the Sun where missions will encounter sufficient ion fluence to exhibit damage to metallic surfaces.

Minow, Joseph I.↗
NASA NTRS2004

Mars Exploration Rover thermal test program overview

In January 2004, two Mars Exploration Rovers (MER) landed on the surface of Mars to begin their mission as robotic geologists. A year prior to these historic landings, both rovers and the spacecraft that delivered them to Mars, were completing a series of environmental tests in facilities at the Jet Propulsion Laboratory. This paper describes the test program undertaken to validate the thermal design and verify the workmanship integrity of both rovers and the spacecraft. The spacecraft, which contained the rover within the aeroshell, were tested in a 7.5 m diameter thermal vacuum chamber. Thermal balance was performed for the near earth (hot case) condition and for the near Mars (cold case) condition. A solar simulator was used to provide the solar boundary condition on the solar array. IR lamps were used to simulate the solar heat load on the aeroshell for the off-sun attitudes experienced by the spacecraft during its cruise to Mars. Each rover was tested separately in a 3.0 m diameter thermal vacuum chamber over conditions simulating the warmest and coldest expected Mars diurnal temperature cycles. The environmental tests were conducted in a quiescent nitrogen atmosphere at a pressure of 8 to 10 Torr. In addition to thermal balance testing, the science instruments on board the rovers were tested successfully in the extreme environmental conditions anticipated for the mission. A solar simulator was not used in these tests.

thermal vacuum test↗
NASA NTRS2003

Thin Film Technology Development for the Powersphere

The Aerospace Corporation, NASA Glenn Research Center, Lockheed-Martin, and ILC Dover over the past two years have been engaged in developing a Multifunctional Inflatable Structure for the Powersphere Concept under contract with NASA (NAS3-01115). The Powersphere concept consists of a relatively large spherical solar array, which would be deployed from a microsatellite. The Powersphere structure and the deployment method was patented by the Aerospace Corporation (U.S. Patent Numbers 6,284,966 B 1 and 6,3 18,675). The work on this project has resulted in a number of technological innovations in the state of the art for integrating flexible thin-film solar cells with flex circuit harness technology and inflatable ultraviolet-light-rigidizable structures. The specific power, specific volume, for the Powersphere are presented in Figures 1 and 2 as a function of solar cell technology and efficiency. The Powersphere will enable microsatellite missions across NASA enterprises and DoD missions by providing ample electric power at an affordable cost. The Powersphere design provides attitude-independent electric power and thermal control for an enclosed microsatellite payload. The design is scalable, robust in high radiation environments and provides sufficient electric power to allow the use of electric propulsion. Electric propulsion enables precise positioning of microsatellites which is required for inspectors that would be deployed to inspect the International Space Station, Space Shuttle or large unmanned spacecraft. The Powersphere allows for efficient launch packaging versus deployed volume as shown in Figure 3.

Simburger, Edward J.↗
NASA NTRS2005

Closed-Cycle Engine Program Used to Study Brayton Power Conversion

One form of power conversion under consideration in NASA Glenn Research Center's Thermal Energy Conversion Branch is the closed-Brayton-cycle engine. In the tens-of-kilowatts to multimegawatt class, the Brayton engine lends itself to potential space nuclear power applications such as electric propulsion or surface power. The Thermal Energy Conversion Branch has most recently concentrated its Brayton studies on electric propulsion for Prometheus. One piece of software used for evaluating such designs over a limited tradeoff space has been the Closed Cycle Engine Program (CCEP). The CCEP originated in the mid-1980s from a Fortran aircraft engine code known as the Navy/NASA Engine Program (NNEP). Components such as a solar collector, heat exchangers, ducting, a pumped-loop radiator, a nuclear heat source, and radial turbomachinery were added to NNEP, transforming it into a high-fidelity design and performance tool for closed-Brayton-cycle power conversion and heat rejection. CCEP was used in the 1990s in conjunction with the Solar Dynamic Ground Test Demonstration conducted at Glenn. Over the past year, updates were made to CCEP to adapt it for an electric propulsion application. The pumped-loop radiator coolant can now be n-heptane, water, or sodium-potassium (NaK); liquid-metal pump design tables were added to accommodate the NaK fluid. For the reactor and shield, a user can now elect to calculate a higher fidelity mass estimate. In addition, helium-xenon working-fluid properties were recalculated and updated.

Johnson, Paul K.↗
NASA NTRS2017

Near Earth Asteroid Solar Sail Engineering Development Unit Test Program

The Near Earth Asteroid (NEA) Scout project is a 30x20x10cm (6U) cubesat reconnaissance mission to investigate a near Earth asteroid utilizing an 86m2 solar sail as the primary propulsion system. This will be the largest solar sail NASA will launch to date. NEA Scout is a secondary payload currently manifested on the maiden voyage of the Space Launch System in 2018. In development of the solar sail subsystem, design challenges were identified and investigated for packaging within a 6U form factor and deployment in cis-lunar space. Analysis furthered understanding of thermal, stress, and dynamics of the stowed system and matured an integrated sail membrane model for deployed flight dynamics. This paper will address design, fabrication, and lessons learned from the NEA Scout solar sail subsystem engineering development unit. From optical properties of the sail material to folding and spooling the single 86m2 sail, the team has developed a robust deployment system for the solar sail. This paper will also address expected and received test results from ascent vent, random vibration, and deployment tests.

Lockett, Tiffany Russell↗