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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 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

An evaluation of some special techniques for nuclear waste disposal in space

A preliminary examination is reported of several special ways for space disposal of nuclear waste material which utilize the radioactive heat in the waste to assist in the propulsion for deep space trajectories. These include use of the wastes in a thermoelectric generator (RTG) which operates an electric propulsion device and a radioisotope - thermal thruster which uses hydrogen or ammonia as the propellant. These propulsive devices are compared to the space tug and the space tug/solar electric propulsion combination for disposal of waste on a solar system escape trajectory. Such comparisons indicate that the waste-RTG approach has considerable potential provided the combined specific mass of the waste container - RTG system does not exceed approximately 150 kg/kw sub e. Several exploratory numerical calculations have been made for high earth orbit and Earth escape destinations.

Mackay, J. S.↗
NASA NTRS1976

Modular thrust subsystem approaches to solar electric propulsion module design

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.

Cake, J. E.↗
NASA NTRS2004

Solar Sail Propulsion Technology Readiness Level Database

The NASA In-Space Propulsion Technology (ISPT) Projects Office has been sponsoring 2 solar sail system design and development hardware demonstration activities over the past 20 months. Able Engineering Company (AEC) of Goleta, CA is leading one team and L Garde, Inc. of Tustin, CA is leading the other team. Component, subsystem and system fabrication and testing has been completed successfully. The goal of these activities is to advance the technology readiness level (TRL) of solar sail propulsion from 3 towards 6 by 2006. These activities will culminate in the deployment and testing of 20-meter solar sail system ground demonstration hardware in the 30 meter diameter thermal-vacuum chamber at NASA Glenn Plum Brook in 2005. This paper will describe the features of a computer database system that documents the results of the solar sail development activities to-date. Illustrations of the hardware components and systems, test results, analytical models, relevant space environment definition and current TRL assessment, as stored and manipulated within the database are presented. This database could serve as a central repository for all data related to the advancement of solar sail technology sponsored by the ISPT, providing an up-to-date assessment of the TRL of this technology. Current plans are to eventually make the database available to the Solar Sail community through the Space Transportation Information Network (STIN).

Adams, Charles L.↗
NASA NTRS2001

The Hybrid Propellant Module (HPM): A New Concept for Space Transfer in the Earth's Neighborhood and Beyond

The safe, affordable and effective transfer of ever-larger payloads and eventually personnel beyond Low Earth Orbit (LEO) is a major challenge facing future commercial development and human exploration of space. Without reusable systems, sustained exploration or large scale development beyond LEO appears to be economically non-viable. However, reusable systems must be capable of both good fuel efficiency and "high utilization of capacity", or else economic costs will remain unacceptably high. Various options exist that can provide high fuel efficiency - for example, Solar Electric Propulsion Systems (SEPS) - but only at the cost of low thrust and concomitant long transit times. Chemical propulsion systems offer the potential for high thrust and short transit times - including both cryogenic and non-cryogenic options - but only at the cost of relatively low specific impulse (Isp). Nuclear thermal propulsion systems offer relatively good thrust-to-weight and Isp - but involve public concerns that may be insurmountable for all except the most-critical of public purposes. Fixed infrastructures have been suggested as one approach to solving this challenge; for example, rotating tether approaches. However, these systems tend to suffer from high initial costs or unacceptable operational constraints. A new concept has been identified - the Hybrid Propellant Module (HPM) - that integrates the best features of both chemical and solar electric transportation architectures. The HPM approach appears to hold promise of solving the issues associated with other approaches, opening a new family of capabilities for future space exploration and development of near-Earth space and beyond. This paper provides a summary overview of the challenge of Earth neighborhood transportation and discusses how various systems concepts might be applied to meet the needs of these architectures. The paper describes a new approach, the HPM, and illustrates the application of the concept for a typical mission concept. The paper concludes with a discussion of needed technologies and a possible timeline for the development and evolution of this class of systems concepts.

Mankins, John C.↗
NASA NTRS2015

Atmospheric Mining in the Outer Solar System: Aerial Vehicle Mission and Design Issues

Atmospheric mining in the outer solar system has been investigated as a means of fuel production for high energy propulsion and power. Fusion fuels such as Helium 3 (3He) and deuterium can be wrested from the atmospheres of Uranus and Neptune and either returned to Earth or used in-situ for energy production. Helium 3 and deuterium were the primary gases of interest with hydrogen being the primary propellant for nuclear thermal solid core and gas core rocket-based atmospheric flight. A series of analyses were undertaken to investigate resource capturing aspects of atmospheric mining in the outer solar system. This included the gas capturing rate, storage options, and different methods of direct use of the captured gases. While capturing 3He, large amounts of hydrogen and 4He are produced. With these two additional gases, the potential for fueling small and large fleets of additional exploration and exploitation vehicles exists. The mining aerospacecraft (ASC) could fly through the outer planet atmospheres, for global weather observations, localized storm or other disturbance investigations, wind speed measurements, polar observations, etc. Analyses of orbital transfer vehicles (OTVs), landers, and in-situ resource utilization (ISRU) mining factories are included. Preliminary observations are presented on near-optimal selections of moon base orbital locations, OTV power levels, and OTV and lander rendezvous points.

orbital mechanics↗
NASA NTRS2008

Space Nuclear Power and Propulsion: Materials Challenges for the 21st Century

The current focus of NASA s space fission effort is Fission Surface Power (FSP). FSP systems could be used to provide power anytime, anywhere on the surface of the Moon or Mars. FSP systems could be used at locations away from the lunar poles or in permanently shaded regions, with no performance penalty. A potential reference 40 kWe option has been devised that is cost-competitive with alternatives while providing more power for less mass. The potential reference system is readily extensible for use on Mars. At Mars the system could be capable of operating through global dust storms and providing year-round power at any Martian latitude. To ensure affordability, the potential near-term, 40 kWe reference concept is designed to use only well established materials and fuels. However, if various materials challenges could be overcome, extremely high performance fission systems could be devised. These include high power, low mass fission surface power systems; in-space systems with high specific power; and high performance nuclear thermal propulsion systems. This tutorial will provide a brief overview of space fission systems and will focus on materials challenges that, if overcome, could help enable advanced exploration and utilization of the solar system.

Houts, Mike↗
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 NASAs exploration goals, a number of projects are developing extensible technologies to support NASAs 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-kW 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.

Electric Propulsion↗
NASA NTRSDate not supplied

Thermal Protection Systems (TPS) for High Velocity Earth Entry Missions

The fastest man-made object to re-enter into Earth’s atmosphere, with an entry velocity of 12.8 km/s, was the Stardust capsule, which returned samples from the comet Wild 2. Sample return missions offer high and long-lasting science yield; as such, they are discussed repeatedly in the latest planetary science decadal survey, titled Origins, Worlds, and Life. An advantage of sample return missions over in situ measurements is the ability to use state-of-the-art instruments that are not limited in power, size, or complexity; conversely, in situ measurements rely on instruments optimized to fit within the limitations of the spacecraft. Additionally, sample return provides the opportunity to revisit samples as analysis techniques improve as well as to further investigate unanticipated or ambiguous results. As launch vehicle and propulsion technologies continue to advance, it is becoming feasible to consider sample return missions from further out in the Solar System beyond Mars orbit. However, returning samples from increasingly distant destinations comes with high entry velocities, and thus requires ever more capable thermal protection systems (TPS). Evaluating the capabilities of existing mature TPS materials for a range of entry conditions and sample return aeroshell configurations is key to establishing the feasibility of future mission proposals. As such, this presentation describes the comprehensive trade study done to evaluate the feasibility of high velocity Earth entry missions with the currently available TPS materials.

Hannah S Alpert↗
NASA NTRSDate not supplied

Thermal Protection Systems (TPS) for High Velocity Earth Entry Missions

The fastest man-made object to re-enter into Earth’s atmosphere, with an entry velocity of 12.8 km/s, was the Stardust capsule, which returned samples from the comet Wild 2. Sample return missions offer high and long-lasting science yield; as such, they are discussed repeatedly in the latest planetary science decadal survey, titled Origins, Worlds, and Life. An advantage of sample return missions over in situ measurements is the ability to use state-of-the-art instruments that are not limited in power, size, or complexity; conversely, in situ measurements rely on instruments optimized to fit within the limitations of the spacecraft. Additionally, sample return provides the opportunity to revisit samples as analysis techniques improve as well as to further investigate unanticipated or ambiguous results. As launch vehicle and propulsion technologies continue to advance, it is becoming feasible to consider sample return missions from further out in the Solar System beyond Mars orbit. However, returning samples from increasingly distant destinations comes with high entry velocities, and thus requires ever more capable thermal protection systems (TPS). Evaluating the capabilities of existing mature TPS materials for a range of entry conditions and sample return aeroshell configurations is key to establishing the feasibility of future mission proposals. As such, this presentation describes the comprehensive trade study done to evaluate the feasibility of high velocity Earth entry missions with the currently available TPS materials.

Hannah S. Alpert↗
NASA NTRS2016

Capabilities of the Environmental Effects Branch at Marshall Space Flight Cente

The Environmental Effects Branch at the Marshall Space Flight Center supports a myriad array of programs for NASA, DoD, and commercial space including human exploration, advanced space propulsion, improving life on Earth, and the study of the Sun, the Earth, and the solar system. The branch provides testing, evaluation, and qualification of materials for use on external spacecraft surfaces and in contamination-sensitive systems. Space environment capabilities include charged particle radiation, ultraviolet radiation, atomic oxygen, impact, plasma, and thermal vacuum, anchored by flight experiments and analysis of returned space hardware. These environmental components can be combined for solar wind or planetary surface environment studies or to evaluate synergistic effects. The Impact Testing Facility allows simulation of impacts ranging from sand and rain to micrometeoroids and orbital debris in order to evaluate materials and components for flight and ground-based systems. The Contamination Control Team is involved in the evaluation of environmentally-friendly replacements for HCFC-225 for use in propulsion oxygen systems, developing cleaning methods for additively manufactured hardware, and reducing risk for the Space Launch System.

Rogers, Jan↗
NASA NTRS2016

Atmospheric Mining in the Outer Solar System: Outer Planet Orbital Transfer and Lander Analyses

Atmospheric mining in the outer solar system has been investigated as a means of fuel production for high energy propulsion and power. Fusion fuels such as Helium 3 (3He) and deuterium can be wrested from the atmospheres of Uranus and Neptune and either returned to Earth or used in-situ for energy production. Helium 3 and deuterium were the primary gases of interest with hydrogen being the primary propellant for nuclear thermal solid core and gas core rocket-based atmospheric flight. A series of analyses were undertaken to investigate resource capturing aspects of atmospheric mining in the outer solar system. This included the gas capturing rate, storage options, and different methods of direct use of the captured gases. While capturing 3He, large amounts of hydrogen and 4He are produced. Analyses of orbital transfer vehicles (OTVs), landers, and the issues with in-situ resource utilization (ISRU) mining factories are included. Preliminary observations are presented on near-optimal selections of moon base orbital locations, OTV power levels, and OTV and lander rendezvous points. For analyses of round trip OTV flights from Uranus to Miranda or Titania, a 10-Megawatt electric (MWe) OTV power level and a 200-metric ton (MT) lander payload were selected based on a relative short OTV trip time and minimization of the number of lander flights. A similar optimum power level is suggested for OTVs flying from low orbit around Neptune to Thalassa or Triton. Several moon base sites at Uranus and Neptune and the OTV requirements to support them are also addressed.

rocket propulsion↗
NASA NTRS2016

Atmospheric Mining in the Outer Solar System: Outer Planet Orbital Transfer and Lander Analyses

Atmospheric mining in the outer solar system has been investigated as a means of fuel production for high energy propulsion and power. Fusion fuels such as Helium 3 (3He) and deuterium can be wrested from the atmospheres of Uranus and Neptune and either returned to Earth or used in-situ for energy production. Helium 3 and deuterium were the primary gases of interest with hydrogen being the primary propellant for nuclear thermal solid core and gas core rocket-based atmospheric flight. A series of analyses were undertaken to investigate resource capturing aspects of atmospheric mining in the outer solar system. This included the gas capturing rate, storage options, and different methods of direct use of the captured gases. While capturing 3He, large amounts of hydrogen and 4He are produced. Analyses of orbital transfer vehicles (OTVs), landers, and the issues with in-situ resource utilization (ISRU) mining factories are included. Preliminary observations are presented on near-optimal selections of moon base orbital locations, OTV power levels, and OTV and lander rendezvous points. For analyses of round trip OTV flights from Uranus to Miranda or Titania, a 10- Megawatt electric (MWe) OTV power level and a 200 metricton (MT) lander payload were selected based on a relative short OTV trip time and minimization of the number of lander flights. A similar optimum power level is suggested for OTVs flying from low orbit around Neptune to Thalassa or Triton. Several moon base sites at Uranus and Neptune and the OTV requirements to support them are also addressed.

lunar mission design↗
NASA NTRS1998

H2OTSTUF: Appropriate Operating Regimes for Magnetohydrodynamic Augmentation

A trade study of magnetohydrodynamic (MHD) augmented propulsion reveals a unique operating regime at lower thrust levels. Substantial mass savings are realized over conventional chemical, solar, and electrical propulsion concepts when MHD augmentation is used to obtain optimal I(sub sp). However, trip times for the most conservative estimates of power plant specific impulse and accelerator efficiency may be prohibitively long. Quasi-one-dimensional calculations show that a solar or nuclear thermal system augmented by MHD can provide competitive performance while utilizing a diverse range of propellants including water, which is available from the Space Shuttle, the Moon, asteroids, and various moons and planets within our solar system. The use of in-situ propellants will reduce costs of space operations as well as enable human exploration of our Solar System. The following conclusions can be drawn from the results of the mission trade study: (1) There exists a maximum thrust or mass flow rate above which MHD augmentation increases the initial mass in low earth orbit (LEO); (2) Mass saving of over 50% can be realized for unique combination of solar/MHD systems; (3) Trip times for systems utilizing current power supply technology may be prohibitively long. Theoretical predictions of MHD performance for in space propulsion systems show that improved efficiencies can reduce trip times to acceptable levels; (4) Long trip times indicative of low thrust systems can be shortened by an increase in the MHD accelerator efficiency or a decrease in the specific mass of the power supply and power processing unit; and (5) As for all propulsion concepts, missions with larger (Delta)v's benefit more from the increased specific impulse resulting from MHD augmentation. Using a quasi-one-dimensional analysis, the required operating conditions for a MHD accelerator to reach acceptable efficiencies are outlined. This analysis shows that substantial non-equilibrium ionization is desirable.

Jones, Jonathan E.↗
NASA NTRS1976

Low-G fluid transfer technology study

Technology gaps and system characteristics critical to cryogenic and noncryogenic in-orbit fluid transfer were identified. Four different supply systems were conceptually designed as space shuttle payloads. These were; (1) space tug supply - LH2, LO2, N2H4, He - linear acceleration for liquid acquisition with supply module and tug separated from shuttle, (2) tug supply using orbiter drag, (3) orbiter supply - N2O4,MMH,He, H2,O2 - surface tension screens, (4) multiple receivers supply 0 solar electric propulsion stage, Hg, diaphragm - HEAO B, HEe, paddle fluid rotation-satellite control section, N2H4, screens. It was found that screens had the best overall potential for low weight and simplicity, however, thermal problems with cryogenics still need final resolution.

Stark, J. A.↗
NASA NTRS1986

The other side of the safety coin

The development, inspection and testing requirements for successful production and launch and safe operation of spaceflight hardware are discussed. Emphasis is placed on paying acute attention to malfunctions, which could be caused by contaminants (particles in docking rings), insufficiently durable materials (Orbiter brakes), etc. Generic and specific problems which occur in propulsion, avionics, mechanical and computer systems and in configuration management, manufacturing and process control efforts are explored. Case histories of deficiencies found in LOX fuel lines, contaminated hydraulic control systems, the Solar Maximum Mission thermal insulation grommets, are summarized. Thorough inspection and testing procedures and design change recording during manufacture of spacecraft components are identified as requisites for successful space missions.

Roth, Gilbert L.↗
NASA NTRS2005

Evaluation of Microbolometer-Based Thermography for Gossamer Space Structures

In August 2003, NASA's In-Space Propulsion Program contracted with our team to develop a prototype on-board Optical Diagnostics System (ODS) for solar sail flight tests. The ODS is intended to monitor sail deployment as well as structural and thermal behavior, and to validate computational models for use in designing future solar sail missions. This paper focuses on the thermography aspects of the ODS. A thermal model was developed to predict local sail temperature variations as a function of sail tilt to the sun, billow depth, and spectral optical properties of front and back sail surfaces. Temperature variations as small as 0.5 C can induce significant thermal strains that compare in magnitude to mechanical strains. These thermally induced strains may result in changes in shape and dynamics. The model also gave insight into the range and sensitivity required for in-flight thermal measurements and supported the development of an ABAQUS-coupled thermo-structural model. The paper also discusses three kinds of tests conducted to 1) determine the optical properties of candidate materials; 2) evaluate uncooled microbolometer-type infrared imagers; and 3) operate a prototype imager with the ODS baseline configuration. (Uncooled bolometers are less sensitive than cooled ones, but may be necessary because of restrictive ODS mass and power limits.) The team measured the spectral properties of several coated polymer samples at various angles of incidence. Two commercially available uncooled microbolometer imagers were compared, and it was found that reliable temperature measurements are feasible for both coated and uncoated sides of typical sail membrane materials.

Miles, Jonathan J.↗
NASA NTRS2012

COMPASS Final Report: Lunar Relay Satellite (LRS)

The Lunar Relay Satellite (LRS) COllaborative Modeling and Parametric Assessment of Space Systems (COMPASS) session was tasked to design a satellite to orbit in an elliptical lunar polar orbit to provide relay communications between lunar South Pole assets and the Earth. The design included a complete master equipment list, power requirement list, configuration design, and brief risk assessment and cost analysis. The LRS is a half-TDRSS sized box spacecraft, which provides communications and navigation relay between lunar outposts (via Lunar Communications Terminals (LCT)) or Sortie parties (with user radios) and large ground antennas on Earth. The LRS consists of a spacecraft containing all the communications and avionics equipment designed by NASA Jet Propulsion Laboratory s (JPL) Team X to perform the relay between lunar-based assets and the Earth. The satellite design is a standard box truss spacecraft design with a thermal control system, 1.7 m solar arrays for 1 kWe power, a 1 m diameter Ka/S band dish which provides relay communications with the LCT, and a Q-band dish for communications to/from the Earth based assets. While JPL's Team X and Goddard Space Flight Center s (GSFC) I M Design Center (IMDC) have completed two other LRS designs, this NASA Glenn Research Center (GRC) COMPASS LRS design sits between them in terms of physical size and capabilities.

Oleson, Steven R.↗