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At least 451 records · Page 25
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Pressurized Lunar Rover

The pressurized lunar rover (PLR) consists of a 7 m long, 3 m diameter cylindrical main vehicle and a trailer which houses the power and heat rejection systems. The main vehicle carries the astronauts, life support systems, navigation and communication systems, directional lighting, cameras, and equipment for exploratory experiments. The PLR shell is constructed of a layered carbon-fiber/foam composite. The rover has six 1.5 m diameter wheels on the main body and two 1.5 m diameter wheels on the trailer. The wheels are constructed of composites and flex to increase traction and shock absorption. The wheels are each attached to a double A-arm aluminum suspension, which allows each wheel 1 m of vertical motion. In conjunction with a 0.75 m ground clearance, the suspension aids the rover in negotiating the uneven lunar terrain. The 15 N-m torque brushless electric motors are mounted with harmonic drive units inside each of the wheels. The rover is steered by electrically varying the speeds of the wheels on either side of the rover. The PLR trailer contains a radiosotope thermoelectric generator providing 6.7 kW. A secondary back-up energy storage system for short-term high-power needs is provided by a bank of batteries. The trailer can be detached to facilitate docking of the main body with the lunar base via an airlock located in the rear of the PLR. The airlock is also used for EVA operation during missions. Life support is a partly regenerative system with air and hygiene water being recycled. A layer of water inside the composite shell surrounds the command center. The water absorbs any damaging radiation, allowing the command center to be used as a safe haven during solar flares. Guidance, navigation, and control are supplied by a strapdown inertial measurement unit that works with the on-board computer. Star mappers provide periodic error correction. The PLR is capable of voice, video, and data transmission. It is equipped with two 5 W X-band transponder, allowing simultaneous transmission and reception. An S-band transponder is used to communicate with the crew during EVA. The PLR has a total mass of 6197 kg. It has a nominal speed of 10 km/hr and a top speed of 18 km/hr. The rover is capable of towing 3 metric tons (in addition to the RTG trailer).

Creel, Kenneth↗
NASA NTRS1992

Small Stirling dynamic isotope power system for robotic space missions

The design of a multihundred-watt Dynamic Isotope Power System (DIPS), based on the U.S. Department of Energy (DOE) General Purpose Heat Source (GPHS) and small (multihundred-watt) free-piston Stirling engine (FPSE), is being pursued as a potential lower cost alternative to radioisotope thermoelectric generators (RTG's). The design is targeted at the power needs of future unmanned deep space and planetary surface exploration missions ranging from scientific probes to Space Exploration Initiative precursor missions. Power level for these missions is less than a kilowatt. The incentive for any dynamic system is that it can save fuel and reduce costs and radiological hazard. Unlike DIPS based on turbomachinery conversion (e.g. Brayton), this small Stirling DIPS can be advantageously scaled to multihundred-watt unit size while preserving size and mass competitiveness with RTG's. Stirling conversion extends the competitive range for dynamic systems down to a few hundred watts--a power level not previously considered for dynamic systems. The challenge for Stirling conversion will be to demonstrate reliability and life similar to RTG experience. Since the competitive potential of FPSE as an isotope converter was first identified, work has focused on feasibility of directly integrating GPHS with the Stirling heater head. Thermal modeling of various radiatively coupled heat source/heater head geometries has been performed using data furnished by the developers of FPSE and GPHS. The analysis indicates that, for the 1050 K heater head configurations considered, GPHS fuel clad temperatures remain within acceptable operating limits. Based on these results, preliminary characterizations of multihundred-watt units have been established.

Bents, D. J.↗
NASA NTRS1992

Design and structural analysis of Mars Rover RTG

The Mars Rover and Sample Return mission's radioisotope thermoelectric generator (RTG) is presently subjected to a structural and mass analysis in view of a reference mission scenario, an illustrative Rover design and Martian activities agenda, and RTG power system requirements and environmental constraints. The modular heat-source stack in the Rover RTG can be held together by axial load springs. The RTGs should be mounted on the Rover with a vertical orientation, in order to avoid the buildup of windborne Martian sand on its heat-rejection surfaces.

Schock, Alfred↗
NASA NTRS1992

Toward the last frontier - A strategy for the evolutionary development of space nuclear power systems

A number of exciting mission opportunities are being considered for the 21st century, including advanced robotic science missions to the outer planets and beyond, human exploration of the Moon and Mars, and advanced space transportation systems. All of these missions will require some form of nuclear power; however, it is clear that current budgetary constraints preclude developing many different types of space nuclear power systems. This paper reviews the specific civil space missions which have been identified, the power levels and lifetimes required, and the technologies available. From this an evolutionary space nuclear power program is developed which builds upon the experience of radioisotope thermoelectric generators, dynamic isotope power systems, and space nuclear reactors. It is strongly suggested that not only does this approach make technical and budgetary sense but that it is consistent with the normal development of new technologies.

Gary L Bennett↗
NASA NTRS1994

Milliwatt Electric Power Sources Would Last For Years

Report discusses design of proposed radioisotope thermoelectric generators, each of which produces about 30 mW of electric power for decades, fits in package about the size of D cell, and would have mass of only 70 g.

Chmielewski, Arthur↗
NASA NTRS1992

Project PARAS: Phased array radio astronomy from space

An orbiting radio telescope is proposed which, when operated in a very long baseline interferometry (VLBI) scheme, would allow higher than currently available angular resolution and dynamic range in the maps and the ability to observe rapidly changing astronomical sources. Using passive phased array technology, the proposed design consists of 656 hexagonal modules forming a 150-m diameter antenna dish. Each observatory module is largely autonomous, having its own photovoltaic power supply and low-noise receiver and processor for phase shifting. The signals received by the modules are channeled via fiber optics to the central control computer in the central bus module. After processing and multiplexing, the data are transmitted to telemetry stations on the ground. The truss frame supporting each observatory panel is a novel hybrid structure consisting of a bottom graphite/epoxy tubular triangle and rigidized inflatable Kevlar tubes connecting the top observatory panel and the bottom triangle. Attitude control and station keeping functions will be performed by a system of momentum wheels in the bus and four propulsion modules located at the compass points on the periphery of the observatory dish. Each propulsion module has four monopropellant thrusters and four hydrazine arcjets, the latter supported by either a photovoltaic array or a radioisotope thermoelectric generator. The total mass of the spacecraft is about 20,500 kg.

Nuss, Kenneth↗
NASA NTRS1992

Pressurized Lunar Rover (PLR)

The objective of this project was to design a manned pressurized lunar rover (PLR) for long-range transportation and for exploration of the lunar surface. The vehicle must be capable of operating on a 14-day mission, traveling within a radius of 500 km during a lunar day or within a 50-km radius during a lunar night. The vehicle must accommodate a nominal crew of four, support two 28-hour EVA's, and in case of emergency, support a crew of six when near the lunar base. A nominal speed of ten km/hr and capability of towing a trailer with a mass of two mt are required. Two preliminary designs have been developed by two independent student teams. The PLR 1 design proposes a seven meter long cylindrical main vehicle and a trailer which houses the power and heat rejection systems. The main vehicle carries the astronauts, life support systems, navigation and communication systems, lighting, robotic arms, tools, and equipment for exploratory experiments. The rover uses a simple mobility system with six wheels on the main vehicle and two on the trailer. The nonpressurized trailer contains a modular radioisotope thermoelectric generator (RTG) supplying 6.5 kW continuous power. A secondary energy storage for short-term peak power needs is provided by a bank of lithium-sulfur dioxide batteries. The life support system is partly a regenerative system with air and hygiene water being recycled. A layer of water inside the composite shell surrounds the command center allowing the center to be used as a safe haven during solar flares. The PLR 1 has a total mass of 6197 kg. It has a top speed of 18 km/hr and is capable of towing three metric tons, in addition to the RTG trailer. The PLR 2 configuration consists of two four-meter diameter, cylindrical hulls which are passively connected by a flexible passageway, resulting in the overall vehicle length of 11 m. The vehicle is driven by eight independently suspended wheels. The dual-cylinder concept allows articulated as well as double Ackermann steering. The primary power of 8 kW is supplied by a dynamic isotope system using a closed Brayton cycle with a xenon-hydrogen mixture as the working fluid. A sodium-sulfur battery serves as the secondary power source. Excess heat produced by the primary power system and other rover systems is rejected by radiators located on the top of the rear cylinder. The total mass of the PLR 2 is 7015 kg. Simplicity and low total weight have been the driving principles behind the design of PLR 1. The overall configuration consists of a 7-m-long, 3-m-diameter cylindrical main vehicle and a two-wheeled trailer. The cylinder of the main body is capped by eight-section, faceted, semi-hemispherical ends. The trailer contains the RTG power source and is not pressurized. The shell of the main body is constructed of a layered carbon fiber/foam/Kevlar sandwich structure. Included in the shell is a layer of water for radiation protection. The layer of water extends from the front of the rover over the crew compartment and creates a safe haven for the crew during a solar flare-up. The carbon fiber provides the majority of the strength and stiffness and the Kevlar provides protection from micrometeoroids. The Kevlar is covered with a gold foil and multi-layer insulation (MLI) to reduce radiation degradation and heat transfer through the wall. A thin thermoplastic layer seals the fiber and provides additional strength.

Creel, Kenneth↗
NASA NTRS1993

An overview of NASA's projected mission requirements for space nuclear systems

The near-term ongoing NASA space nuclear program is the Cassini mission to Saturn which will use three radioisotope thermoelectric generators (RTGs). Beyond that NASA is studying a Pluto Fast Flyby mission which will challenge the space power community to produce a low-mass RTG. Another candidate RTG mission is the Mars Environmental Survey (MESUR) mission to emplace a number of probes on the surface of Mars to obtain a more global survey of the planet than was accomplished with the two Viking Landers. Looking toward the 21st century, there are a number of exciting planetary missions, such as the Jupiter Grand Tour, Outer Planet Orbiters/Probes, comet/asteroid rendezvous/sample return, which are enabled or greatly enhanced by nuclear reactor power coupled with electric propulsion.

Bennett, Gary L.↗
NASA NTRS1994

Project ECHO: Electronic Communications from Halo Orbit

The design of a communications relay to provide constant access between the Earth and the far side of the Moon is presented. Placement of the relay in a halo orbit about the L2 Earth-Moon Lagrange point allows the satellite to maintain constant simultaneous communication between Earth and scientific payloads on the far side of the Moon. The requirements of NASA's Discovery-class missions adopted and modified for this design are: total project cost should not exceed $150 million excluding launch costs, launch must be provided by Delta-class vehicle, and the satellite should maintain an operational lifetime of 10 to 15 years. The spacecraft will follow a transfer trajectory to the L2 point, after launch by a Delta II 7925 vehicle in 1999. Low-level thrust is used for injection into a stationkeeping-free halo orbit once the spacecraft reaches the L2 point. The shape of this halo orbit is highly elliptical with the maximum excursion from the L2 point being 35000 km. A spun section and despun section connected through a bearing and power transfer assembly (BAPTA) compose the structure of the spacecraft. Communications equipment is placed on the despun section to provide for a stationary dual parabolic offset-feed array antenna system. The dual system is necessary to provide communications coverage during portions of maximum excursion on the halo orbit. Transmissions to the NASA Deep Space Network 34 m antenna include six channels (color video, two voice, scientific data from lunar payloads, satellite housekeeping and telemetry and uplinked commands) using the S- and X-bands. Four radioisotope thermoelectric generators (RTG's) provide a total of 1360 W to power onboard systems and any two of the four Hughes 13 cm ion thrusters at once. Output of the ion thrusters is approximately 17.8 mN each with xenon as the propellant. Presence of torques generated by solar pressure on the antenna dish require the addition of a 'skirt' extending from the spun section of the satellite for balance. Total mass of the satellite is approximately 900 kg at a cost of $130 million FY99.

Borrelli, Jason↗
NASA NTRS1995

Thermal barrier coatings application in diesel engines

Commercial use of thermal barrier coatings in diesel engines began in the mid 70's by Dr,. Ingard Kvernes at the Central Institute for Industrial Research in Oslo, Norway. Dr. Kvernes attributed attack on diesel engine valves and piston crowns encountered in marine diesel engines in Norwegian ships as hot-corrosion attributed to a reduced quality of residual fuel. His solution was to coat these components to reduce metal temperature below the threshold of aggressive hot-corrosion and also to provide protection. The Department of Energy has supported thermal barrier coating development for diesel engine applications. In the Clean Diesel - 50 Percent Efficient (CD-50) engine for the year 2000, thermal barrier coatings will be used on piston crowns and possibly other components. The primary purpose of the thermal barrier coatings will be to reduce thermal fatigue as the engine peak cylinder pressure will nearly be doubled. As the coatings result in higher available energy in the exhaust gas, efficiency gains are achieved through use of this energy by turbochargers, turbocompounding or thermoelectric generators.

Fairbanks, J. W.↗
NASA NTRS1995

Development of advanced Si and GaAs solar cells for interplanetary missions

The deep space and planetary exploration project have been acquiring more and more importance and some of them are now well established both in ESA and NASA programs. This paper presents the possibility to utilize both silicon and gallium arsenide solar cells as spacecraft primary power source for missions far from the Sun, in order to overcome the drawbacks related to the utilisation of radioisotope thermoelectric generators - such as cost, safety and social acceptance. The development of solar cells for low illumination intensity and low temperature (LILT) applications is carried out in Europe by ASE (Germany) and CISE (Italy) in the frame of an ESA programme, aimed to provide the photovoltaic generators for ROSETTA: the cometary material investigation mission scheduled for launch in 2003. The LILT cells development and testing objectives are therefore focused on the following requirements: insolation intensity as low as 0.03 Solar Constant, low temperature down to -150 C and solar flare proton environment. At this stage of development, after the completion of the technology verification tests, it has been demonstrated that suitable technologies are available for the qualification of both silicon and gallium arsenide cells and both candidates have shown conversion efficiencies over 25% at an illumination of 0.03 SC and a temperature of -150 C. In particular, when measured at those LILT conditions, the newly developed 'Hl-ETA/NR-LILT' silicon solar cells have reached a conversion efficiency of 26.3%, that is the highest value ever measured on a single junction solar cell. A large quantity of both 'Hl-ETA/NR-LILT' silicon and 'GaAs/Ge-LILT' solar cells are presently under fabrication and they will be submitted to a qualification test plan, including radiation exposure, in order to verify their applicability with respect to the mission requirements. The availability of two valid options will minimize the risk for the very ambitious scientific project. The paper describes how the technical achievements have been possible with Si and GaAs LILT solar cells (including a comparison between measured and modelled l-V characteristics) and it presents the technology verification tests results.

Strobl, G.↗
NASA NTRS1995

Thermal barrier coatings application in diesel engines

Commercial use of thermal barrier coatings in diesel engines began in the mid 70's by Dr. Ingard Kvernes at the Central Institute for Industrial Research in Oslo, Norway. Dr. Kvernes attributed attack on diesel engine valves and piston crowns encountered in marine diesel engines in Norwegian ships as hot-corrosion attributed to a reduced quality of residual fuel. His solution was to coat these components to reduce metal temperature below the threshold of aggressive hot-corrosion and also provide protection. Roy Kamo introduced thermal barrier coatings in his 'Adiabatic Diesel Engine' in the late 70's. Kamo's concept was to eliminate the engine block water cooling system and reduce heat losses. Roy reported significant performance improvements in his thermally insulated engine at the SAE Congress in 1982. Kamo's work stimulates major programs with insulated engines, particularly in Europe. Most of the major diesel engine manufacturers conducted some level of test with insulated combustion chamber components. They initially ran into increased fuel consumption. The German engine consortium had Prof. Woschni of the Technical Institute in Munich. Woschni conducted testing with pistons with air gaps to provide the insulation effects. Woschni indicated the hot walls of the insulated engine created a major increase in heat transfer he refers to as 'convection vive.' Woschni's work was a major factor in the abrupt curtailment of insulated diesel engine work in continental Europe. Ricardo in the UK suggested that combustion should be reoptimized for the hot-wall effects of the insulated combustion chamber and showed under a narrow range of conditions fuel economy could be improved. The Department of Energy has supported thermal barrier coating development for diesel engine applications. In the Clean Diesel - 50 Percent Efficient (CD-50) engine for the year 2000, thermal barrier coatings will be used on piston crowns and possibly other components. The primary purpose of the thermal barrier coatings will be to reduce thermal fatigue as the engine peak cylinder pressure will nearly be doubled. As the coatings result in higher available energy in the exhaust gas, efficiency gains are achieved through use of this energy by turbochargers, turbocompounding or thermoelectric generators.

Fairbanks, J. W.↗
NASA NTRS1997

Acoustic Testing of the Cassini Spacecraft and Titan IV Payload Fairing: Results - Part 2

A Cassini spacecraft simulator in a full scale 60 foot high Titan 4 payload fairing with various acoustic blanket designs and configurations was recently tested in a large reverberant acoustic chamber. A first part companion paper provides the test configuration details and other background information. This paper addresses the results obtained from this test program. Emphasis will be on the effects of the new blanket designs on reducing the payload fairing's internal acoustics and the vibration response of the spacecraft's Radioisotope Thermoelectric Generators. Other results discussed include: the effect of blankets on fairing vibration, the effect of partial blanket coverage on acoustics and vibration and the effect of tuned vibration absorbers.

Hughes, William O.↗
NASA NTRS1997

Acoustic Testing of the Cassini Spacecraft and Titan 4 Payload Fairing: Introduction and Test Configuration - Part 1

NASA Lewis Research Center recently led a multi-organizational acoustic test program. This testing consisted of acoustically exciting a Cassini spacecraft simulator in a full scale 60 foot high Titan 4 payload fairing with various acoustic blanket designs and configurations in a large reverberant acoustic chamber. The primary purpose of this test program was to measure the fairing's internal acoustics and spacecraft vibration, especially the Radioisotope Thermoelectric Generators (RTG) vibration, and to quantify the mitigation efforts in reducing these levels. Key to this reduction effort was the utilization of new acoustic blanket designs. This paper will provide the background and rationale for performing this test program, state the test program's primary and secondary objectives and describe the test matrix, hardware and instrumentation. A second part companion paper will provide the test results and data analysis.

Hughes, William O.↗
NASA NTRS1997

Environmental Impact Statement for the Cassini Mission

This Final Supplemental Environmental Impact Statement (FSEIS) to the 1995 Cassini mission Environmental Impact Statement (EIS) focuses on information recently made available from updated mission safety analyses. This information is pertinent to the consequence and risk analyses of potential accidents during the launch and cruise phases of the mission that were addressed in the EIS. The type of accidents evaluated are those which could potentially result in a release of plutonium dioxide from the three Radioisotope Thermoelectric Generators (RTGS) and the up to 129 Radioisotope Heater Units (RHUS) onboard the Cassini spacecraft. The RTGs use the heat of decay of plutonium dioxide to generate electric power for the spacecraft and instruments. The RHUs, each of which contains a small amount of plutonium dioxide, provide heat for controlling the thermal environment of the spacecraft and several of its instruments. The planned Cassini mission is an international cooperative effort of the National Aeronautics and Space Administration (NASA), the European Space Agency (ESA), and the Italian Space Agency (ASI) to conduct a 4-year scientific exploration of the planet Saturn, its atmosphere, moons, rings, and magnetosphere.

Source record↗
NASA NTRS1989

RTG Safety Tests

The primary objective of STS-34 was to launch Galileo on its trip to Jupiter. The Galileo spacecraft contains two Radioisotope Thermoelectric Generators (RTG), which contains plutonium. This videotape shows and the accompanying material explains the tests that the RTG containment vessel has been subjected to, and the results of the tests. The videotape shows the trajectory of the Galileo spacecraft, a cutaway view of an RTG, the Plutonium-238 fuel capsule, and seven of the tests on the RTG.

Source record↗
NASA NTRS2000

A Deep Space Power System Option Based on Synergistic Power Conversion Technologies

Deep space science missions have typically used radioisotope thermoelectric generator (RTG) power systems. The RTG power system has proven itself to be a rugged and highly reliable power system over many missions, however the thermal-to-electric conversion technology used was approximately 5% efficient. While the relatively low efficiency has some benefits in terms of system integration, there are compelling reasons why a more efficient conversion system should be pursued. The cost savings alone that are available as a result of the reduced isotope inventory are significant. The Advanced Radioisotope Power System (ARPS) project was established to fulfill this goal. Although it was not part of the ARPS project, Stirling conversion technology is being demonstrated with a low level of funding by both NASA and DOE. A power system with Stirling convertors. although intended for use with an isotope heat source. can be combined with other advanced technologies to provide a novel power system for deep space missions. An inflatable primary concentrator would be used in combination with a refractive secondary concentrator (RSC) as the heat source to power the system. The inflatable technology as a structure has made great progress for a variety of potential applications such as communications reflectors, radiators and solar arrays. The RSC has been pursued for use in solar thermal propulsion applications, and it's unique properties allow some advantageous system trades to be made. The power system proposed would completely eliminate the isotope heat source and could potentially provide power for science missions to planets as distant as Uranus. This paper will present the background and developmental status of the technologies and will then describe the power system being proposed.

Schreiber, Jeffrey G.↗
NASA NTRS2000

Technology Assessment of Doe's 55-we Stirling Technology Demonstrator Convector (TDC)

The Department of Energy (DOE), Germantown, Maryland and the NASA Glenn Research Center (GRC), Cleveland, Ohio are developing a Stirling Convertor for an advanced radioisotope power system as a potential power source for spacecraft on-board electric power for NASA deep space science missions. The Stirling Convertor is being evaluated as an alternative high efficiency power source to replace Radioisotope Thermoelectric Generators (RTGs). Stirling Technology Company (STC), Kennewick, Washington, is developing the highly efficient, long life 55-We free-piston Stirling Convertor known as the Technology Demonstrator Convertor (TDC) under contract to DOE. GRC provides Stirling technology expertise under a Space Act Agreement with the DOE. Lockheed Martin Astronautics (LMA), Valley Forge, Pennsylvania is the current power system integrator for the Advanced Radioisotope Power System (ARPS) Project for the DOE. JPL is responsible for the Outer Planets/Solar Probe Project for NASA.

Furlong, Richard↗