Non-Nuclear Exploration of the Solar System Study

Abstract

Advances in solar array, electric propulsion (EP), and power beaming technologies will very likely enable future missions to the ice giants (Uranus and Neptune) by spacecraft that are completely solar powered. Between 1959 and 2001 the power from solar arrays on missions in Earth orbit have increased by five orders of magnitude (from 1 W to >100,000 W). Simultaneously, the use of solar power has been extended to ever larger distances from the Sun. Three solar powered missions out to solar ranges of just beyond 5 au have now been developed and flown (Rosetta, Juno, and Lucy). At these distances, the solar insolation is roughly 25 times lower than that at 1 au. Solar-powered spacecraft to Saturn, where the solar insolation is 100 times lower than that at 1 au have recently been proposed to NASA. A solar-powered mission to Uranus would have to function where solar insolation is only 4 times lower than it is at Saturn. For Neptune, it would be 9 times lower than at Saturn. Three improvements in solar array technology are required to make this feasible. First, solar cells have to be able to function in the low-intensity, low-temperature (LILT) environment at Uranus and Neptune. Specially designed triple-junction solar cells have been tested at JPL under LILT conditions equivalent to the environment at 30 au have shown excellent performance (high efficiency and high fill factor). Second, the size of deployable solar arrays has to increase by one to two orders of magnitude relative to the current state of the art. Third, the areal density of solar arrays has to be reduced by an order of magnitude. This will most likely be accomplished through a combination of new thin-film solar cell technology like the Perovskite cells under development worldwide and the development of new deployable solar array structures such as those under investigation for solar-powered satellites. The solar insolation at Uranus and Neptune is 400 to 900 times lower, respectively, than it is at 1 au. To provide sufficient power for a spacecraft at these destinations requires very large solar arrays. To be practical, such arrays will necessarily need to be very lightweight with a minimum structure mass. The gentle, low-thrust nature of electric propulsion is a good match for such solar arrays since the continuous acceleration of order 10⁻⁵ g will not drive increases in structure mass. In addition, the availability of large amounts of power provided by these arrays between 1 au and 5 to 10 au enables the design of low-thrust trajectories to the ice giants with attractive flight times. Existing ion propulsion technologies, such as NASA's NEXT ion propulsion system enable flight times of conventionally sized spacecraft (≥1000 kg, not including the solar array) to Uranus of less than 10 years with reasonable propellant masses. For example, a conventionally sized spacecraft with a 150-kg complement of instruments could be delivered to Uranus orbit with a vehicle that has two 60-m x 60-m solar array wings and an areal density of 100 g/m² in a flight time of less than 10 years. The same-sized vehicle could be delivered to orbit around Neptune in a flight time of less than 18 years if the solar array areal density is reduced to 50 g/m². In both cases, larger payload masses can be delivered in similar flight times by increasing the size of the solar array wings. A 440-kg payload mass could be delivered to Neptune orbit in less than 17 years by increasing the solar array size to 70 m x 70 m. Smaller spacecraft may offer nearer-term opportunities. For example, coupling large, lightweight solar arrays with low-power ion propulsion systems (maximum input power of ~3 kW) can deliver net spacecraft masses to Uranus orbit of several hundred kilograms in flight times of less than 10 years. The only new development for such missions would be solar array wings with dimensions of 30 m x 30 m to 60 m x 60 m with an areal density of 100 g/m². The same EP system could deliver net spacecraft masses of 300 to 500 kg (inclusive of the payload, but not including the solar array, xenon tank, and xenon mass margin, which are tracked separately) to Neptune orbit with flight times of around 15 years with a 60-m x 60-m solar array that has an areal density of 50 g/m2. Such an array would provide roughly 2.4 kW at Neptune. The development of directed energy (DE) systems could potentially provide hundreds of watts of power continuously to landed assets from solar-powered orbiting spacecraft. Using a DE system to convert electrical power to light on the orbiting spacecraft and a photo-converter system on the landed asset to convert the DE light back to electricity is effectively a "photonic extension cord." For the DE side, a series of lasers in an array is used to beam power to distant landed assets over distances of hundreds to thousands of kilometers. The landed assets use high efficiency photovoltaics tuned to the laser frequency to convert the laser power back to electrical power. Thermal power from the photon power not converted to electricity may, in some applications, also be useful to the landed asset. State-of-the-art directed-energy systems are solid state, efficient (~50%), low mass (~1 kg/kW_(optical)), and long lifetime (~10⁵ hrs). This technology is improving rapidly, driven by the photonic revolution along with consumer and industrial demand. It is likely even possible to beam power to multiple stationary and even moving targets using unique optical beacons from each target. The technologies required for non-nuclear exploration of the solar system would also enable or enhance a wide variety of other missions of national interest. The large, ultra-light solar arrays combined with a state-of-the-art electric propulsion system would make possible the orbital exploration of Pluto, as well as a tour of its large moon Charon and smaller moons, with a reasonable mass margin and could potentially eliminate the need for RTGs for this mission. Large, ultra-light solar arrays and state-of-the-art electric propulsion systems could enable missions to chase down and encounter long-period comets and potentially even interstellar objects. This combination of technologies could enable solar electric propulsion (SEP) mission architectures farther out in the solar system, including: a Kuiper belt tour, centaur tour, or maybe even a 'Grand Tour' of the gas and ice giants without need for the most optimal planetary alignment, as with the Voyager missions. Sample returns are the next frontier in planetary exploration. The SEP and solar array technologies discussed here would facilitate sample return from a wide range of bodies, including possibly Ceres, Mars, Enceladus, Titan, Triton, and maybe even Mercury. Beaming power from a large, ultra-light solar array in orbit to a landed asset could enable non-nuclear surface exploration of the ice giant's moons. In the nearer term, directed energy systems could deliver power to the surface of the Moon or Mars. Large, lighter solar arrays could facilitate lower-risk human missions to Mars using very high-power solar electric propulsion systems in mission architectures that don't require rendezvousing with pre-deployed assets for the return trip. Finally, large, ultra-light solar arrays could power an Arecibo-like radar in space, to enhance characterization of potentially hazardous objects. Such high-power solar arrays combined with ion-beam deflection would be greatly enhance the nations's planetary defense capabilities.

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