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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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Phobos Surveyor Mission concept provides an innovative low cost, highly reliable approach to exploring the inner solar system 1/16/2013 3 Dual manifest launch. Use only flight proven, well characterize commercial off-the-shelf components. Flexible mission architecture allows for a slew of unique measurements.
The Ice and Fire Preproject defined the first mission set of the Outer Planets/Solar Probe program consisting of the Europa Orbiter, Pluto-Kuiper Express, and Solar Probe missions.
Electric propulsion ion engine systems using solar power source for Mars and Jupiter exploratory unmanned spacecraft is adaptable to existing launch vehicles
Outer-planet mission and systems analyses are performed using three next generation solar-electric ion thruster models. The impact of variations in thruster model, flight time, launch vehicle, propulsion and power systems characteristics is investigated. All presented trajectories have a single Venus gravity assist and maximize the delivered mass to Saturn or Neptune. The effect of revolution ratio - the ratio of Venusian orbital period to the flight time between launch and flyby dates - is also discussed.
Feasibility design study of solar powered ion propulsion system for interplanetary spacecraft
The comet Encke at its apparition in 1980 is investigated as the target for a solar electric slow flyby mission having a flight time of approximately two years. Mission trade off studies are performed for two modes: namely the direct earth to Encke mode and the Venus gravity assist mode. Baseline missions for each mode are selected which flyby Encke at 4 km/s 30 days before perihelion. The direct mission baseline is investigated in detail for trajectory parameters related to spacecraft and science requirements. The Venus flyby mode can deliver the same payload with a total power requirement about 20% less than the direct mode. The cost is a longer flight time and a Venus flyby at about 500 km above the surface. One asteroid can be encountered at very close range on either mission. The asteroids available and the extra propellant required are determined.
Centaur launch vehicle for Jupiter and Saturn orbiter missions using solar electric propelled spacecraft
Photographic mapping of Martian surface with electrically propelled solar-powered photographic spacecraft placed into near polar orbit
During the past decade, a number of advances have occurred in solar cell and array technology. These advances have lead to performance improvement for both conventional space arrays and for advanced technology arrays. Performance enhancements have occurred in power density, specific power, and environmental capability. Both state-of-the-art and advanced development cells and array technology are discussed. Present technology will include rigid, rollout, and foldout flexible substrate designs, with silicon and GaAs solar cells. The use of concentrator array systems is also discussed based on both DOD and NASA efforts. The benefits of advanced lightweight array technology, for both near term and far term utilization, and of advanced high efficiency, thin, radiation resistant cells is examined. This includes gallium arsenide on germaniun substrates, indium phosphide, and thin film devices such as copper indium diselenide.
Recent NASA interest in utilizing solar electronic propulsion (SEP) technology to transfer payloads, e.g. from low-Earth orbit (LEO) to higher energy geostationary-Earth orbit (GEO) or to Earth escape, has necessitated the development of high fidelity SEP vehicle models and simulations. These models and simulations need to be capable of capturing vehicle dynamics and sub-system interactions experienced during the transfer trajectories which are typically accomplished with continuous-burn (potentially interrupted by solar eclipse), long duration "spiral out" maneuvers taking several months or more to complete. This paper presents details of an integrated simulation approach achieved by combining a high fidelity vehicle simulation code with a detailed solar array model. The combined simulation tool gives researchers the functionality to study the integrated effects of various vehicle sub-systems (e.g. vehicle guidance, navigation and control (GN&C), electric propulsion system (EP)) with time varying power production. Results from a simulation model of a vehicle with a 50 kW class SEP system using the integrated tool are presented and compared to the results from another simulation model employing a 50 kW end-of-life (EOL) fixed power level assumption. These models simulate a vehicle under three degree of freedom dynamics (i.e. translational dynamics only) and include the effects of a targeting guidance algorithm (providing a "near optimal" transfer) during a LEO to near Earth escape (C (sub 3) = −2.0 km (sup 2) / sec (sup −2) spiral trajectory. The presented results include the impact of the fully integrated, time-varying solar array model (e.g. cumulative array degradation from traversing the Van Allen belts, impact of solar eclipses on the vehicle and the related temperature responses in the solar arrays due to operating in the Earth's thermal environment, high fidelity array power module, etc.); these are used to assess the impact on vehicle performance (i.e. propellant consumption) and transit times.
Design approaches are described and evaluated for a mercury electron-bombardment ion thruster array. Such an array might be used on a solar electric interplanetary spacecraft that obtains electrical energy from large solar panels. Thruster array designs are described and evaluated as they would apply to an Encke Flyby mission. Besides several well known approaches, a new concept utilizing individual two-axis gimbal actuators on each thruster is described and shown to have many structural and thermal advantages.
Solar sails have long been known to be an attractive method of propulsion in the inner solar system if the areal density of the overall spacecraft (S/C) could be reduced to approx.10 g/sq m. It has also long been recognized that the figure (precise shape) of useful solar sails needs to be reasonably good, so that the reflected light goes mostly in the desired direction. If one could make large reflective surfaces with reasonable figure at an areal density of approx.10 g/sq m, then several other attractive options emerge. One is to use such sails as solar concentrators for solar-electric propulsion. Current flight solar arrays have a specific output of approx. 100W/kg at 1 Astronomical Unit (AU) from the sun, and near-term advances promise to significantly increase this figure. A S/C with an areal density of 10 g/sq m could accelerate up to 29 km/s per year as a solar sail at 1 AU. Using the same sail as a concentrator at 30 AU, the same spacecraft could have up to approx. 45 W of electric power per kg of total S/C mass available for electric propulsion (EP). With an EP system that is 50% power-efficient, exhausting 10% of the initial S/C mass per year as propellant, the exhaust velocity is approx. 119 km/s and the acceleration is approx. 12 km/s per year. This hybrid thus opens attractive options for missions to the outer solar system, including sample-return missions. If solar-thermal propulsion were perfected, it would offer an attractive intermediate between solar sailing in the inner solar system and solar electric propulsion for the outer solar system. In the example above, both the solar sail and solar electric systems don't have a specific impulse that is near-optimal for the mission. Solar thermal propulsion, with an exhaust velocity of the order of 10 km/s, is better matched to many solar system exploration missions. This paper derives the basic relationships between these three propulsion options and gives examples of missions that might be enabled by such hybrids.
A computer-controlled attitude and articulation control system is proposed for a spacecraft of extremely soft structure that uses two 8-kW solar arrays with an area of nearly 100 sq m each to power six electric propulsion thrusters. The purpose of the articulation control is to point the scientific instruments, to orient the antenna toward the earth, and to reorient the thrust vector as required for trajectory control. The attitude control system provides control both during the low-thrust phase of the mission and during long-term spinning modes.
This paper defines an advanced SEP concept based upon demonstrated ion thruster technology improvements and projected improvements in solar array technology. The ASEP concept utilizes an 'integrated' rather than a 'stage' approach and assumes 30 cm Hg ion thrusters capable of operation near 7 kWe each. The mission performance using the ASEP system is presented for three missions: (1) Saturn orbiter, (2) comet rendezvous and (3) comet sample return. The ASEP system provides significant mission performance improvements. The simpler, fewer thruster ASEP system design may provide cost as well as performance benefits for the missions examined when compared to ballistic or SEP stage mission options.
A revised user's manual for the computer program MAPSEP is presented. Major changes from the interplanetary version of MAPSEP are summarized. The changes are intended to provide a basic capability to analyze anticipated solar electric missions, and a foundation for future more complex, modifications. For Vol. III, N75-16589.
Established procedures of linear, quadratic, Gaussian optimal estimation and control are developed and interpreted for their application to the problem of attitude control of spacecraft with dynamically significant elastic appendages. Results are presented both in general terms and for specific application to a solar electric spacecraft. Comparisons are made between alternative coordinate systems, and a realistic range of design parameters is considered. For single axis control, system evaluation is accomplished by simulation of a fifteenth-order spacecraft plant with alternative second-order, fourth-order, and sixth-order constant gain estimators. Results point out the importance of modeling errors.
In the current National Aeronautics and Space Administration (NASA) budget plan, there exists a new proposed program aimed at opening up the outer solar system to new robotic science missions.
The propulsion module comprises six to eight 30-cm thruster and power processing units, a mercury propellant storage and distribution system, a solar array ranging in power from 18 to 25 kW, and the thermal and structure systems required to support the thrust and power subsystems. Launch and on-orbit configurations are presented for both modular approaches. The propulsion module satisfies the thermal design requirements of a multimission set including: Mercury, Saturn, and Jupiter orbiters, a 1-AU solar observatory, and comet and asteroid rendezvous. A detailed mass breakdown and a mass equation relating the total mass to the number of thrusters and solar array power requirement is given for both approaches.