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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 235 records · Page 13
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NASA NTRS1986

The Pegasus drive - A multi-megawatt nuclear electric propulsion system

Pegasus, a power generating system for use in space, is multimegawatt power system that would enable missions of almost any conceivable duration and scope. The Pegassus Drive is the coupling of this nuclear electric power system with a 6 MWe MPD thruster. The power system has a maximum power output of 8.5 MWe. The MPD thruster requires 6 MWe to provide spacecraft propulsion and 1.5 MWe are available for mission-specific tasks and experiments. The balance of power generated is used to operate the power system. The size and mass limitations of the STS are of prime consideration in the design of this system to allow the collapsed system to be placed in lower earth orbit by two Shuttle missions. The main system (19,120 kg) employing a shadow shield would require one launch and the balance of the four-pi shield (27,830 kg) would occupy the second launch. Development of this power system could be completed by the mid 1990's and the system available near the turn of the century.

Coomes, E. P.↗
NASA NTRS2018

Magnetour: Surfing Planetary Systems on Electromagnetic and Multi-Body Gravity Fields

In this NIAC Phase One study, we propose a new mission concept, named Magnetour, to facilitate the exploration of outer planet systems and address both power and propulsion challenges. Our approach would enable a single spacecraft to orbit and travel between multiple moons of an outer planet, with no propellant required. Our approach would enable a single spacecraft to orbit and travel between multiple moons of an outer planet, with no propellant nor onboard power source required. To achieve this free-lunch _Grand Tour', we exploit the unexplored combination of magnetic and multi-body gravitational fields of planetary systems, with a unique focus on using a bare tether for power and propulsion. The main objective of the study is to develop this conceptually novel mission architecture, explore its design space, and investigate its feasibility and applicability to enhance the exploration of planetary systems within a 10-year timeframe. Propellantless propulsion technology offers enormous potential to transform the way NASA conducts outer planet missions. We hope to demonstrate that our free-lunch tour concept can replace heavy, costly, traditional chemical-based missions and can open up a new variety of trajectories around outer planets. Leveraging the powerful magnetic and multi-body gravity fields of planetary systems to travel freely among planetary moons would allow for long-term missions and provide unique scientific capabilities and flagship-class science for a fraction of the mass and cost of traditional concepts. New mission design techniques are needed to fully exploit the potential of this new concept.This final report contains the results and findings of the Phase One study, and is organized as follows. First, an overview of the Magnetour mission concept is presented. Then, the research methodology adopted for this Phase One study is described, followed by a brief outline of the main findings and their correspondence with the original Phase One task plan. Next, an overview of the environment of outer planets is provided, including magnetosphere, radiation belt and planetary moons. Then performance of electrodynamic tethers is assessed, as well as other electromagnetic systems. A method to exploit multi-body dynamics is given next. These analyses allow us to carry out a Jovian mission design to gain insight in the benefits of Magnetour. In addition, a spacecraft configuration is presented that fully incorporates the tether in the design. Finally technology roadmap considerations are discussed.

Gravity Fields↗
NASA NTRS1989

Advanced solar-propelled cargo spacecraft for Mars missions

At the University of Washington, three concepts for an unmanned, solar powered, cargo spacecraft for Mars-support missions have been investigated. These spacecraft are designed to carry a 50,000 kg payload from a low Earth orbit to a low Mars orbit. Each design uses a distinctly different propulsion system: a solar radiation absorption (SRA) system, a solar-pumped laser (SPL) system, and a solar powered mangetoplasmadynamic (MPD) arc system. The SRA directly converts solar energy to thermal energy in the propellant through a novel process developed at the University of Washington. A solar concentrator focuses sunlight into an absorption chamber. A mixture of hydrogen and potassium vapor absorbs the incident radiation and is heated to approximately 3700 K. The hot propellant gas exhausts through a nozzle to produce thrust. The SRA has an I(sub sp) of approximately 1000 sec and produces a thrust of 2940 N using two thrust chambers. In the SPL system, a pair of solar-pumped, multi-megawatt, CO2 lasers in sun-synchronous Earth orbit converts solar energy to laser energy. The laser beams are transmitted to the spacecraft via laser relay satellites. The laser energy heats the hydrogen propellant through a plasma breakdown process in the center of an absorption chamber. Propellant flowing through the chamber, heated by the plasma core, expands through a nozzle to produce thrust. The SPL has an I(sub sp) of 1285 sec and produces a thrust of 1200 N using two thrust chambers. The MPD system uses indium phosphide solar cells to convert sunlight to electricity, which powers the propulsion system. In this system, the argon propellant is ionized and electromagnetically accelerated by a magnetoplasmadynamic arc to produce thrust. The MPD spacecraft has an I(sub sp) of 2490 sec and produces a thrust of 100 N. Various orbital transfer options are examined for these concepts. In the SRA system, the mother ship transfers the payload into a very high Earth orbit and a small auxiliary propulsion system boosts the payload into a Hohmann transfer to Mars. The SPL spacecraft releases the payload as the spacecraft passes by Mars. Both the SRA-powered spacecraft and the SPL-powered spacecraft return to Earth for subsequent missions. The MPD-propelled spacecraft, however, remains at Mars as an orbiting space station. A patched conic approximation was used to determine a heliocentric interplanetary transfer orbit for the MPD propelled spacecraft. All three solar-powered spacecraft use an aerobrake procedure to place the payload into a low Mars parking orbit. The payload delivery times range from 160 days to 873 days (2.39 years).

Auziasdeturenne, J.↗
NASA NTRS1985

Reactor power system deployment and startup

This paper addresses issues that should receive further examination in the near-term as concept selection for development of a U.S. space reactor power system is approached. The issues include: the economics, practicality and system reliability associated with transfer of nuclear spacecraft from low earth shuttle orbits to operational orbits, via chemical propulsion versus nuclear electric propulsion; possible astronaut supervised reactor and nuclear electric propulsion startup in low altitude Shuttle orbit; potential deployment methods for nuclear powered spacecraft from Shuttle; the general public safety of low altitude startup and nuclear safe and disposal orbits; the question of preferred reactor power level; and the question of frozen versus molten alkali metal coolant during launch and deployment. These issues must be considered now because they impact the SP-100 concept selection, power level selection, weight and size limits, use of deployable radiators, reliability requirements, and economics, as well as the degree of need for and the urgency of developing space reactor power systems.

Wetch, J. R.↗
NASA NTRS1963

THE ELECTRON-BOMBARDMENT ION ROCKET

Proposed electron bombardment ion rocket, discussing ion chamber, magnetic field, cathode, accelerator and neutralizer in terms of spacecraft power requirements

PROPULSION CONFIGURATION↗
NASA NTRS2017

Osiris-REx Spacecraft Current Status and Forward Plans

The NASA New Frontiers OSIRIS-REx spacecraft executed a flawless launch on September 8, 2016 to begin its 23-month journey to near-Earth asteroid (101955). The primary objective of the OSIRIS-REx mission is to collect and return to Earth a pristine sample of regolith from the asteroid surface. The sampling event will occur after a two-year period of remote sensing that will ensure a high probability of successful sampling of a region on the asteroid surface having high science value and within well-defined geological context. The OSIRIS-REx instrument payload includes three high-resolution cameras (OCAMS), a visible and near-infrared spectrometer (OVIRS), a thermal imaging spectrometer (OTES), an X-ray imaging spectrometer (REXIS), and a laser altimeter (OLA). As the spacecraft follows its nominal outbound-cruise trajectory, the propulsion, power, communications, and science instruments have undergone basic functional tests, with no major issues. Outbound cruise science investigations include a search for Earth Trojan asteroids as the spacecraft approaches the Sun-Earth L4 Lagrangian point in February 2017. Additional instrument checkouts and calibrations will be carried out during the Earth gravity assist maneuver in September 2017. During the Earth-moon flyby, visual and spectral images will be acquired to validate instrument command sequences planned for Bennu remote sensing. The asteroid Bennu remote sensing campaign will yield high resolution maps of the temperature and thermal inertia, distributions of major minerals and concentrations of organic matter across the asteroid surface. A high resolution 3d shape model including local surface slopes and a high-resolution gravity field will also be determined. Together, these data will be used to generate four separate maps that will be used to select the sampling site(s). The Safety map will identify hazardous and safe operational regions on the asteroid surface. The Deliverability map will quantify the accuracy with which the navigation team can deliver the spacecraft to and from specific sites on the asteroid surface. The Sampleability map quantifies the regolith properties, providing an estimation of how much material would be sampled at different points on the surface. The final Science Value map synthesizes the chemical, mineralogical, and geological, observations to identify the areas of the asteroid surface with the highest science value. Here, priority is given to organic, water-rich regions that have been minimally altered by surface processes. Asteroid surface samples will be acquired with a touch-and-go sample acquisition system (TAGSAM) that uses high purity pressurized N2 gas to mobilize regolith into a stainless steel canister. Although the mission requirement is to collect at least 60 g of material, tests of the TAGSAM routinely exceeded 300 g of simulant in micro-gravity tests. After acquiring the sample, the spacecraft will depart Bennu in 2021 to begin its return journey, with the sample return capsule landing at the Utah Test and Training Range on September 23, 2023. The OSIRIS-REx science team will carry out a series of detailed chemical, mineralogical, isotopic, and spectral studies that will be used to determine the origin and history of Bennu and to relate high spatial resolution sample studies to the global geological context from remote sensing. The outline of the sample analysis plan is described in a companion abstract.

Messenger, Scott↗
NASA NTRS2015

Testing of an Arcjet Thruster with Capability of Direct-Drive Operation

Electric thrusters typically require a power processing unit (PPU) to convert the spacecraft provided power to the voltage-current that a thruster needs for operation. Testing has been initiated to study whether an arcjet thruster can be operated directly with the power produced by solar arrays without any additional conversion. Elimination of the PPU significantly reduces system-level complexity of the propulsion system, and lowers developmental cost and risk. The work aims to identify and address technical questions related to power conditioning and noise suppression in the system and heating of the thruster in long-duration operation. The apparatus under investigation has a target power level from 400-1,000 W. However, the proposed direct-drive arcjet is potentially a highly scalable concept, applicable to solar-electric spacecraft with up to 100's of kW and beyond. A direct-drive electric propulsion system would be comprised of a thruster that operates with the power supplied directly from the power source (typically solar arrays) with no further power conditioning needed between those two components. Arcjet thrusters are electric propulsion devices, with the power supplied as a high current at low voltage; of all the different types of electric thruster, they are best suited for direct drive from solar arrays. One advantage of an arcjet over Hall or gridded ion thrusters is that for comparable power the arcjet is a much smaller device and can provide more thrust and orders of magnitude higher thrust density (approximately 1-10 N/sq m), albeit at lower I(sub sp) (approximately 800-1000 s). In addition, arcjets are capable of operating on a wide range of propellant options, having been demonstrated on H2, ammonia, N2, Ar, Kr, Xe, while present SOA Hall and ion thrusters are primarily limited to Xe propellant. Direct-drive is often discussed in terms of Hall thrusters, but they require 250-300 V for operation, which is difficult even with high-voltage solar arrays. The arcjet requires under 100 V, which is more in-line with what is easily possible with a solar array. Direct-drive of an electric propulsion system confers the advantage of reducing or eliminating the power processing unit (PPU) that is typically needed to convert the spacecraft-provided power to the voltage and current needed for thruster operation. Since the PPU is typically the most expensive piece of an electric thruster system, from both a fabrication and qualification standpoint, its elimination offers the potential for major reductions in system cost and risk. The design of the arcjet built for this effort was based on previous low power (1 kW class) arcjets. It has a precision machined 99.95% pure tungsten anode which also serves as the nozzle. The anode constrictor region is 1 mm (0.040-in) diameter and 1 mm (0.040-in) long. The cathode is a tungsten welding electrode doped with LaO2; its tip was precision ground to a 30◦ angle ending in a blunt end. The two electrodes are separated by a boron-nitride insulator which also serves as the propellant injection manifold; it ends in six small holes which introduce the propellant gas in the diverging section of the nozzle, directly adjacent to the cathode. The electrodes and insulator are housed in a stainless-steel outer-body, with a Macor insulator at the mid-plane to provide thermal isolation between the front and back halves of the device. The gas seals were made using Grafoil gaskets. Figure 1A shows the assembled thruster in the vacuum chamber; figure 1B shows the thruster in operation on argon at a flow rate of 676 sccm (20 mg/s). Initial testing was conducted in a 3.5-ft diameter vacuum chamber; the ultimate pressure reached during quasi-steady operation of the thruster was about 330 millitorr. The thruster was powered with a high-current, 0-100A, 15 kW power supply. The discharge was initiated with a high-voltage (approximately 10 kV) spark initiator that was isolated from the supply by a stack of diodes. The testing indicated that an operating point exists within the I-V characteristics that is compatible with direct-drive solar-electric operation; for a flow rate of 20 mg/s (argon) the arc could be sustained at a voltage of about 20 V and a current of 25 A (500W).

Martin, Adam K.↗
NASA NTRS1995

Nuclear Thermal Rocket/Vehicle Design Options for Future NASA Missions to the Moon and Mars

The nuclear thermal rocket (NTR) provides a unique propulsion capability to planners/designers of future human exploration missions to the Moon and Mars. In addition to its high specific impulse (approximately 850-1000 s) and engine thrust-to-weight ratio (approximately 3-10), the NTR can also be configured as a 'dual mode' system capable of generating electrical power for spacecraft environmental systems, communications, and enhanced stage operations (e.g., refrigeration for long-term liquid hydrogen storage). At present the Nuclear Propulsion Office (NPO) is examining a variety of mission applications for the NTR ranging from an expendable, single-burn, trans-lunar injection (TLI) stage for NASA's First Lunar Outpost (FLO) mission to all propulsive, multiburn, NTR-powered spacecraft supporting a 'split cargo-piloted sprint' Mars mission architecture. Each application results in a particular set of requirements in areas such as the number of engines and their respective thrust levels, restart capability, fuel operating temperature and lifetime, cryofluid storage, and stage size. Two solid core NTR concepts are examined -- one based on NERVA (Nuclear Engine for Rocket Vehicle Application) derivative reactor (NDR) technology, and a second concept which utilizes a ternary carbide 'twisted ribbon' fuel form developed by the Commonwealth of Independent States (CIS). The NDR and CIS concepts have an established technology database involving significant nuclear testing at or near representative operating conditions. Integrated systems and mission studies indicate that clusters of two to four 15 to 25 klbf NDR or CIS engines are sufficient for most of the lunar and Mars mission scenarios currently under consideration. This paper provides descriptions and performance characteristics for the NDR and CIS concepts, summarizes NASA's First Lunar Outpost and Mars mission scenarios, and describes characteristics for representative cargo and piloted vehicles compatible with a reference 240 t-class heavy lift launch vehicle (HLLV) and smaller 120 t HLLV option. Attractive performance characteristics and high-leverage technologies associated with both the engine and stage are identified, and supporting parametric sensitivity data is provided. The potential for commonality of engine and stage components to satisfy a broad range of lunar and Mars missions is also discussed.

Borowski, Stanley K.↗
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 NTRS1980

The Voyager spacecraft /James Watt International Gold Medal Lecture/

The Voyager Project background is reviewed with emphasis on selected features of the Voyager spacecraft. Investigations by the Thermo-electric Outer Planets Spacecraft Project are discussed, including trajectories, design requirements, and the development of a Self Test and Repair computer, and a Computer Accessed Telemetry System. The design and configuration of the spacecraft are described, including long range communications, attitude control, solar independent power, sequencing and control data handling, and spacecraft propulsion. The development program, maintained by JPL, experienced a variety of problems such as design deficiencies, and process control and manufacturing problems. Finally, the spacecraft encounter with Jupiter is discussed, and expectations for the Saturn encounter are expressed.

Heacock, R. L.↗
NASA NTRS1982

Space exploration: The interstellar goal and Titan demonstration

Automated interstellar space exploration is reviewed. The Titan demonstration mission is discussed. Remote sensing and automated modeling are considered. Nuclear electric propulsion, main orbiting spacecraft, lander/rover, subsatellites, atmospheric probes, powered air vehicles, and a surface science network comprise mission component concepts. Machine, intelligence in space exploration is discussed.

Source record↗
NASA NTRS2002

Shedding Light on Solar Power

Glenn Research Center sponsored an SBIR contract with ENTECH, in which the company worked to mold its successful terrestrial concentrator technology into applications that would generate solar power for space missions. ENTECH's first application made use of small, dome-shaped Fresnel lenses to direct sunlight onto high- efficiency photovoltaic cells. After some key adjustments, the mini- dome lens array was flown as part of the U.S. Air Force/NASA Photovoltaic Array Space Power Plus Diagnostics (PASP Plus) flight experiment in 1994. Due to their three-dimensional shape, the mini- dome lenses entailed construction by a batch molding process, which is naturally more costly than a continuous process. To overcome this disadvantage and meet the requirement for precise solar pointing in two axes, ENTECH started developing solar concentrator arrays for space using a line-focus lens that can be mass-produced by a continuous process. This new technology, named Solar Concentrator Array with Refractive Linear Element Technology (SCARLET), was created with support from Glenn and the Ballistic Missile Defense Organization, and was used to power the NASA/Jet Propulsion Laboratory Deep Space 1 spacecraft.

Source record↗