Paul Schmitz

Venus Rover Design Study

The surface of Venus is a hostile environment, with a surface temperature averaging 452°C, and atmospheric pressure of 92 bars of carbon dioxide. Nevertheless, exploration of the surface of Venus is of scientific interest. Technologies currently being developed at NASA bring the operation of robots on Venus into the range of feasibility. These include high-temperature electronics; radioisotope power systems that operate at Venus temperatures; Stirling-based power and cooling systems to keep mission components within temperature constraints; high-temperature, corrosion-resistant materials; and hightemperature sensors, motors, actuators, and bearings. This paper presents the results of a design study for a Venus rover, with the goal of surface exploration capability that is comparable to that of Mars rover missions. The most critical portion of the design is the power and cooling system required for operation at Venus, and this analysis will comprise most of the work presented. The thermal design requires operation at an external temperature as high as 500°C. To minimize external heat flow into the electronics enclosure (and also to provide maximum structural strength against external pressure) the electronics enclosure is assumed to be spherical shell. A radioisotope Stirling Duplex engine provides electrical power and 2-stage cooling. In order to minimize heat leaks, the number of penetrations to the thermal enclosure was minimized. Optical instruments were assumed to operate through sapphire windows, and all the components that could be operated in the high-temperature Venus environment were located outside the enclosure. Total rover design mass (including cruise and EDL system) is 872 kg without growth, and 1059 kg with mass growth allowance included; this is easily within the launch capability of an expendable launch vehicle. There are no apparent showstoppers to the design of a rover capable of operation on the Venus surface, although some of the technologies will still need development to bring them to flight readiness.

Radioisotope Electric Propulsion Centaur Orbiter Spacecraft Design Overview

Radioisotope electric propulsion (REP) has been shown in past studies to enable missions to outerplanetary bodies including the orbiting of Centaur asteroids. Key to the feasibility for REP missions are long life, low power electric propulsion (EP) devices, low mass radioisotope power systems (RPS) and light spacecraft (S/C) components. In order to determine what are the key parameters for EP devices to perform these REP missions a design study was completed to design an REP S/C to orbit a Centaur in a New Frontiers cost cap. The design shows that an orbiter using several long lived (approximately 200 kg Xenon throughput), low power (approximately 700 W) Hall thrusters teamed with six (150 W each) Advanced Stirling Radioisotope Generators (ASRG) can deliver 60 kg of science instruments to a Centaur in 10 yr within the New Frontiers cost cap. Optimal specific impulses for the Hall thrusters were found to be around 2000 sec with thruster efficiencies over 40%. Not only can the REP S/C enable orbiting a Centaur (when compared to an all chemical mission only capable of flybys) but the additional power from the REP system can be reused to enhance science and simplify communications.

Solar vs. Fission Surface Power for Mars

A multi-discipline team of experts from the National Aeronautics and Space Administration (NASA) developed Mars surface power system point design solutions for two conceptual missions. The primary goal of this study was to compare the relative merits of solar- versus fission-powered versions of each surface mission. First, the team compared three different solar power options against a fission power system concept for a sub-scale, uncrewed demonstration mission. The 4.5 meter (m) diameter pathfinder landers primary mission would be to demonstrate Mars entry, descent, and landing techniques. Once on the Martian surface, the landers In Situ Resource Utilization (ISRU) payload would demonstrate liquid oxygen propellant production using atmospheric resources. For the purpose of this exercise, location was assumed to be at the Martian equator. The three solar concepts considered included a system that only operated during daylight hours (at roughly half the daily propellant production rate of a round-the-clock fission design), a battery-augmented system that operated through the night (matching the fission concepts propellant production rate), and a system that operated only during daylight, but at a higher rate (again, matching the fission concepts propellant production rate). Including 30% mass growth allowance, total payload masses for the three solar concepts ranged from 1,116 to 2,396 kg, versus the 2,686 kg fission power scheme. However, solar power masses are expected to approach or exceed the fission payload mass at landing sites further from the equator, making landing site selection a key driver in the final power system decision. The team also noted that detailed reliability analysis should be performed on daytime-only solar power schemes to assess potential issues with frequent ISRU system on/off cycling. Next, the team developed a solar-powered point design solution for a conceptual four-crew, 500-day surface mission consisting of up to four landers per crewed expedition mission. Unlike the demonstration mission, a lengthy power outage due to the global dust storms that are known to occur on Mars would pose a safety hazard to a crewed mission. A similar fission versus solar power trade study performed by NASA in 2007 concluded that fission power was more reliable-with a much lower mass penalty-than solar power for this application. However, recent advances in solar cell and energy storage technologies and changes in operational assumptions prompted NASA to revisit the analysis. For the purpose of this exercise a particular landing site at Jezero Crater, located at 18o north latitude, was assumed. A fission power system consisting of four each 10 kW Kilopower fission reactors was compared to a distributed network of Orion-derived Ultraflex solar arrays and Lithium ion batteries mounted on every lander. The team found that a solar power system mass of about 9,800 kg would provide the 22 kilowatts (kW) keep-alive power needed to survive a dust storm lasting up to 120-days at average optical depth of 5, and 35 kW peak power for normal operations under clear skies. Although this is less than half the mass estimated during the 2007 work (which assumed latitudes up to 30o) it is still more than the 7,000 kg mass of the fission system which provides full power regardless of dust storm conditions.

MASER: A Mars meteorology and seismology mini-network mission concept enabled by Milliwatt-RPS

A design reference mission (MASER - Meteorology and Seismology Enabled by Radioisotopes) for a Mars mini-network is presented. Four hard landers using parachutes and crushable impact attenuators would be deployed in the polar plains north of the Tharsis bulge to perform seismic and meteorological measurements throughout a Martian year (including the dark winter). Operation throughout the polar winter is only possible through the use of a power subsystem that would rely upon six Radioisotope Heater Units (RHUs), providing ~240mWe.

Dynamic Radioisotope Power System development for space exploration

Dynamic power conversion offers the potential to produce Radioisotope Power Systems (RPS) that generate higher power outputs and utilize the available heat source plutonium fuel more efficiently than Radioisotope Thermoelectric Generators. Additionally, dynamic systems offer the potential of producing generators with significantly reduced power degradation over the course of deep space missions so that more power would be available at the end of the mission, when it is needed most for both powering science instruments and transmitting the resulting data. The development of dynamic generators involves addressing technical issues not typically associated with traditional thermoelectric generators. Developing long-life, robust, and reliable dynamic conversion technology is challenging yet essential to building a suitable flight-ready generator. Considerations include working within existing hardware-handling infrastructure, where possible, so that development costs can be kept low, and integrating dynamic generators into spacecraft, which may be more complex than integration of static thermoelectric systems. Methods of interfacing to and controlling a dynamic generator must also be considered, and new potential failure modes must be taken into account. This paper will address some of the key issues of dynamic RPS design, development, and adaption.