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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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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.
The technologies which will permit sub-0.1 arcsec pointing accuracies on spacecraft in the 1990s are examined, along with the accuracies required and the current state of the art. Of particular interest are multi-mission spacecraft. Pointing accuracy can only be obtained by integrating the instrument (telescope) as part of the spacecraft, minimizing disturbances and using reaction wheels for pointing. The pointer could be isolated from complex spacecraft disturbances by soft mechanical mounts, e.g., inflatible tethers, guy-wire suspension and fluidic pointing systems. All design options are being explored for the Space Station, Earth Observing System, Co-orbiting platform and GEO platform spacecraft, and for near-term planetary spacecraft which will employ nuclear electric propulsion.
A collection of scaling equations, weight statements, scaling factors, etc., useful for doing conceptual designs of spacecraft are given. Rules of thumb and methods of calculating quantities of interest are provided. Basic relationships for conventional, and several non-conventional, propulsion systems (nuclear, solar electric and solar thermal) are included. The equations and other data were taken from a number of sources and are not at all consistent with each other in level of detail or method, but provide useful references for early estimation purposes.
The present evaluation of highly detailed advanced propulsion system design concepts for the Mars Rover Sample Return Mission proceeded by comparing a baseline chemical propulsion option with both storable and cryogenic advanced chemical propulsion alternatives and solar- and nuclear-based electric propulsion OTVs. Substantial launch mass reductions and commensurate payload mass increases were obtainable with both advanced chemical and electric propulsion cycles.
The design and performance of a spacecraft employing arcjet nuclear electric propulsion, suitable for use in the SP-100 Space Reactor Power System (SRPS) Flight Experiment, are outlined. The vehicle design is based on a 93 kW(e) ammonia arcjet system operating at an experimentally measured specific impulse of 1031 s and an efficiency of 42.3 percent. The arcjet/gimbal assemblies, power conditioning subsystem, propellant feed system, propulsion system thermal control, spacecraft diagnostic instrumentation, and the telemetry requirements are described. A 100 kW(e) SRPS is assumed. The spacecraft mass is baselined at 5675 kg excluding the propellant and propellant feed system. Four mission scenarios are described which are capable of demonstrating the full capability of the SRPS. The missions considered include spacecraft deployment to possible surveillance platform orbits, a spacecraft storage mission, and an orbit raising round trip corresponding to possible orbit transfer vehicle (OTV) missions.
The design and performance of a spacecraft employing arcjet nuclear electric propulsion, suitable for use in the SP-100 Space Reactor Power System (SRPS) Flight Experiment, are outlined. The vehicle design is based on a 93 kWe ammonia arcjet system operating at an experimentally-measured specific impulse of 1030 s and an efficiency of 42 percent. The arcjet/gimbal assemblies, power conditioning subsystem, propellant feed system, propulsion system thermal control, spacecraft diagnostic instrumentation, and the telemetry requirements are described. A 100 kWe SRPS is assumed. The total spacecraft mass is baselined at 5675 kg excluding the propellant and propellant feed system. Four mission scenarios are described which are capable of demonstrating the full capability of the SRPS. The missions considered include spacecraft deployment to possible surveillance platform orbits, a spacecraft storage mission and an orbit raising round trip corresponding to possible orbit transfer vehicle missions. Launches from Kennedy Space Center using the Titan IV expendable launch vehicle are assumed.
The design and performance of a spacecraft employing arcjet nuclear-electric propulsion, suitable for use in the SP-100 space-reactor power-system (SRPS) flight experiment, are outlined. The vehicle design is based on a 93-kWe ammonia arcjet system operating at an experimentally measured specific impulse of 1031 sec and an efficiency of 42.3 percent. The arcjet/gimbal assemblies, power conditioning subsystem, propellant feed system, propulsion system thermal control, spacecraft diagnostic instrumentation, and the telemetry requirements are described. A 100-kWe SRPS is assumed. The spacecraft mass is baselined at 5675 kg, excluding the propellant and propellant feed system. Four mission scenarios are described which are capable of demonstrating the full capability of the SRPS.
The documentation of the Trajectory Generation and System Characterization Model for the Cislunar Low-Thrust Spacecraft is presented in Technical and User's Manuals. The system characteristics and trajectories of low thrust nuclear electric propulsion spacecraft can be generated through the use of multiple system technology models coupled with a high fidelity trajectory generation routine. The Earth to Moon trajectories utilize near Earth orbital plane alignment, midcourse control dependent upon the spacecraft's Jacobian constant, and capture to target orbit utilizing velocity matching algorithms. The trajectory generation is performed in a perturbed two-body equinoctial formulation and the restricted three-body formulation. A single control is determined by the user for the interactive midcourse portion of the trajectory. The full spacecraft system characteristics and trajectory are provided as output.
Space reactor power systems with characteristics ideal for advanced spacecraft systems applications are discussed. These characteristics are: high power-to-weight ratio (15 to 33 W/kg); high volume density (high ballistic coefficient); no preferential orientation in orbit; long operational life; high reliability; and total launch and operational safety. These characteristics allow the use of electric propulsion to raise spacecraft from low earth parking orbits to operational orbits, greatly increasing the useful orbit payload for a given launch vehicle by eliminating the need for a separation injection stage. A proposed demonstration mission is described.
The progress and results are summarized for mission/system requirements database; mission analysis; GN and C (Guidance, Navigation, and Control), aeroheating, Mars landing; radiation protection; aerobrake mass analysis; Shuttle-Z, TMIS (Trans-Mars Injection Stage); Long Duration Habitat Trade Study; evolutionary lunar and Mars options; NTR (Nuclear Thermal Rocket); NEP (Nuclear Electric Propulsion) update; SEP (Solar Electric Propulsion) update; orbital and space-based requirements; technology; piloted rover; programmatic task; and evolutionary and innovative architecture.
The Space Eploration Initiative launched by the Bush Administration in May, 1990, will intensively explore alternative mission architectures involving the use of the moon as a staging post for manned exploration of Mars. The advanced technologies to be analyzed by the 'Pathfinder' program for prospective use in the alternative missions thus far defined encompass aerobraking, space-based chemical engines, cryogenic fluid systems, space nuclear power, nuclear-thermal and nuclear-electric propulsion, high-rate communications, and planetary photonics. Extensive use of robotics is anticipated in both the lunar and Martian components of the missions being studied.
It is pointed out that future exploration of the moon and Mars will mandate developments in many areas of technology. In particular, major advances will be required in planet surface power systems. Critical nuclear technology challenges that can enable strategic self-sufficiency, acceptable operational costs, and cost-effective space transportation goals for NASA exploration missions have been identified. Critical technologies for surface power systems include stationary and mobile nuclear reactor and radioisotope heat sources coupled to static and dynamic power conversion devices. These technologies can provide dramatic reductions in mass, leading to operational and transportation cost savings. Critical technologies for space transportation systems include nuclear thermal rocket and nuclear electric propulsion options, which present compelling concepts for significantly reducing mass, cost, or travel time required for Earth-Mars transport.
The manned exploration of Mars will require the development of one or more enabling technologies. Employing the go as you pay philosophy of the Advisory Committee on the Future of the U.S. Space Program, the timelines for the early flights will be strongly influenced by our ability to fashion innovative, productive missions using existing or near-term technologies to the maximum extent possible. In this paper, the potential benefits of nuclear electric propulsion (NEP) employing SP-100 reactor technology for unmanned Mars cargo missions are investigated. Results suggest that only moderate growth of current technology is necessary to deliver useful cargo payloads with reasonable transfer times.
Space System Architecture Code (SSAC) program performs sensitivity studies on operating parameters of space missions. SSAC examines effects of variations in sizes of crews of piloted missions, chemical vs. nuclear electric propulsion, and specific impulse. Written in FORTRAN 77.
The main topics covered include: (1) the Space Exploration Initiative (SEI) context; (2) critical issues of MPD Thruster design; and (3) the Magnetoplasmadynamic (MPD) Thruster Program at JPL. Under the section on the SEI context the nuclear electric propulsion system and some electric thruster options are addressed. The critical issues of MPD Thruster development deal with the requirements, status, and approach taken. The following areas are covered with respect to the MPD Thruster Program at JPL: (1) the radiation-cooled MPD thruster; (2) the High-Current Cathode Test Facility; (3) thruster component thermal modeling; and (4) alkali metal propellant studies.
The development of the wire core system for Nuclear Thermal Propulsion (NTP) that took place from 1963 to 1965 is discussed. A wire core consists of a fuel wire with spacer wires. It's an annular flow core having a central control rod. There are actually four of these, with beryllium solid reflectors on both ends and all the way around. Much of the information on the concept is given in viewgraph form. Viewgraphs are presented on design details of the wire core, the engine design, engine weight vs. thrust, a technique used to fabricate the wire fuel element, and axial temperature distribution.
Future Mars exploration missions, both robotic and piloted, may utilize Earth to Mars transfer trajectories that are significantly different from one another, depending upon the type of mission being flown and the time period during which the flight takes place. The use of new or emerging technologies for future missions to Mars, such as aerobraking and nuclear rocket propulsion, may yield navigation requirements that are much more stringent than those of past robotic missions, and are very difficult to meet for some trajectories. This article explores the interdependencies between the properties of direct Earth to Mars trajectories and the Mars approach navigation accuracy that can be achieved using different radio metric data types, such as ranging measurements between an approaching spacecraft and Mars orbiting relay satellites, or Earth based measurements such as coherent Doppler and very long baseline interferometry. The trajectory characteristics affecting navigation performance are identified, and the variations in accuracy that might be experienced over the range of different Mars approach trajectories are discussed. The results predict that three sigma periapsis altitude navigation uncertainties of 2 to 10 km can be achieved when a Mars orbiting satellite is used as a navigation aid.
Nuclear Electric Propulsion (NEP) mission performance is strongly affected by system performance. Power and propulsion system specific mass, specific impulse, and efficiency all combine to determine the performance limits for a given mission. Thruster technology determines the specific impulse and efficiency of the system. The effects of these parameters on the mission performance of NEP systems relative to other concepts was analyzed to give guidance to thruster development goals for a range of missions: Lunar Cargo, Mars Cargo, and Mars Piloted. Mission sensitivities to system parameters are discussed, and technology requirements are identified for each mission.