Thomas W. Kerslake

Analysis of Thermal Energy Storage Material With Change-of-Phase Volumetric Effects

NASAs Space Station Freedom proposed hybrid power system includes photovoltaic arrays with nickel hydrogen batteries for energy storage and solar dynamic collectors driving Brayton heat engines with change-of-phase thermal energy storage (TES) devices. A TES device is comprised of multiple metallic, annular canisters which contain a eutectic composition LiF-CaF[sub 2] phase change material (PCM) that melts at 1,040 K. A moderately sophisticated LiF-CaF[sub 2] PCM computer model is being developed in two stages considering first one-dimensional and then two-dimensional canister geometries. One-dimensional model results indicate that the void has a marked effect on the phase change process due to PCM displacement and dynamic void heat transfer resistance. Equally influential are the effects of different boundary conditions and liquid PCM free convection. For the second stage, successful numerical techniques used in the one-dimensional phase change model are extended to a two-dimensional (r, z) PCM containment canister model. A prototypical PCM containment canister is analyzed and the results are discussed.

Options for the human exploration of Mars using Solar Electric propulsion

Solar Electric propulsion (SEP) is examined as a candidate transportation option for human missions to Mars. Focus is given to an Earth-escape staging concept. This concept uses a SEP system to transfer from low earth orbit (LEO) to a high-energy elliptical parking orbit (HEEPO) and a chemical propulsion system to transfer from the HEEPO to a hyperbolic escape trajectory. LEO to Earth escape performance of these combined transportation systems is comparable to that of a nuclear thermal rocket (NTR). As a result, a mass efficient non-nuclear transportation architecture with fast, 180 day, Earth-to-Mars piloted transit times is enabled.

Benefits of Power and Propulsion Technology for a Piloted Electric Vehicle to an Asteroid

Abstract NASA’s goal for human spaceflight is to expand permanent human presence beyond low Earth orbit (LEO). NASA is identifying potential missions and technologies needed to achieve this goal. Mission options include crewed destinations to LEO and the International Space Station; high Earth orbit and geosynchronous orbit; cis-lunar space, lunar orbit, and the surface of the Moon; near-Earth objects; and the moons of Mars, Mars orbit, and the surface of Mars. NASA generated a series of design reference missions to drive out required functions and capabilities for these destinations, focusing first on a piloted mission to a near-Earth asteroid. One conclusion from this exercise was that a solar electric propulsion stage could reduce mission cost by reducing the required number of heavy lift launches and could increase mission reliability by providing a robust architecture for the long-duration crewed mission. Similarly, solar electric vehicles were identified as critical for missions to Mars, including orbiting Mars, landing on its surface, and visiting its moons. This paper describes the parameterized assessment of power and propulsion technologies for a piloted solar electric vehicle to a near-Earth asteroid. The objective of the assessment was to determine technology drivers to advance the stateof the art of electric propulsion systems for human exploration. Sensitivity analyses on the performance characteristics of the propulsion and power systems were done to determine potential system-level impacts of improved technology. Starting with a “reasonable vehicle configuration” bounded by an assumed launch date, we introduced technology improvements to determine the system-level benefits (if any) that those technologies might provide. The results of this assessment are discussed and recommendations for future work are described.

Thin-Film Photovoltaic Solar Array Parametric Assessment

This paper summarizes a study that had the objective to develop a model and parametrically determine the circumstances for which lightweight thin-film photovoltaic solar arrays would be more beneficial, in terms of mass and cost, than arrays using highefficiency crystalline solar cells. Previous studies considering arrays with near-term thin-film technology for Earth orbiting applications are briefly reviewed. The present study uses a parametric approach that evaluated the performance of lightweight thin-film arrays with cell efficiencies ranging from 5% to 20%. The model developed for this study is described in some detail. Similar mass and cost trends for each array option were found across eight missions of various power levels in locations ranging from Venus to Jupiter. The results for one specific mission, a main belt asteroid tour, indicate that only moderate thin-film cell efficiency (-12%) is necessary to match the mass of arrays using crystalline cells with much greater efficiency (35% multi-junction GaAs based and 20% thin-silicon). Regarding cost, a 12% efficient thin-film array is projected to cost about half as much as a 4junction GaAs array. While efficiency improvements beyond 12% did not significantly further improve the mass and cost benefits for thin-film arrays, higher efficiency will be needed to mitigate the spacecraftlevel impacts associated with large deployed array areas. A low-temperatur e approach to depositing thinfilm cells on lightweight, flexible plastic substrates is briefly described. The paper concludes with the observation that with the characteristics assumed for this study, ultra-lightweight arrays using efficient, thinfilm cells on flexible substrates may become a leading alternative for a wide variety of space missions.

Feasibility of Large High-Powered Solar Electric Propulsion Vehicles: Issues and Solutions

Human exploration beyond low earth orbit will require the use of enabling technologies that are efficient, affordable, and reliable. Solar electric propulsion (SEP) has been proposed by NASA’s Human Exploration Framework Team (HEFT) as an option to achieve human exploration missions to Near Earth Objects (NEOs) because of its favorable mass efficiency as compared to traditional chemical systems. This paper describes the unique challenges and technology hurdles associated with developing a large high-power SEP vehicle. A subsystem level breakdown of factors contributing to the feasibility of SEP as a platform for future exploration missions to NEOs is presented including overall mission feasibility, trip time variables, propellant management issues, solar array power generation, array structure issues, and other areas that warrant investment in additional technology or engineering development.

Modeling Cyclic Phase Change and Energy Storage in Solar Heat Receivers

Numerical results pertaining to cyclic melting and freezing of an encapsulated phase-change material (PCM), integrated into a solar heat receiver, have been reported. The cyclic nature of the present problem is relevant to latent heat thermal energy storage systems used to power solar Brayton engines in space. Specifically, a physical and numerical model of the solar heat receiver component of NASA Lewis Research Centers ground test demonstration (GTD) system was developed and results compared with available experimental data. Multiconjugate effects such as the convective flow of a low Prandtl number fluid, conduction in the PCM, containment tube, and working fluid conduit were accounted for in the model. An ideal-body thermal radiation model was also included to quantify reradiative effects inside the receiver along with losses through the aperture and receiver shell. A high-temperature eutectic mixture of LiF-CaF 2 was used as the PCM and a mixture of He/Xe was used as the working fluid. A modified version of the computer code HOTTube was used to generate results for comparisons with GTD experimental data in both subcooled and two-phase regimes. While qualitative trends were in close agreement for the balanced-orbit modes, excellent quantitative agreement was observed for steady-state modes.

Solar power system analyses for electric propulsion missions

Solar electric propulsion (SEP) mission architectures are applicable to a wide range of NASA missions including human Mars exploration and robotic exploration of the outer planets. In this paper, we discuss the conceptual design and detailed performance analysis of an SEP stage electric power system (EPS). EPS performance, mass and area predictions are compared for several PV array technologies. Based on these studies, an EPS design for a 1-MW class, Human Mars Mission SEP stage was developed with a reasonable mass, 9.4 metric tons, and feasible deployed array area, 5800 m/sup 2/. An EPS was also designed for the Europa Mapper spacecraft and had a mass of 151 kg and a deployed array area of 106 m/sup 2/.

Validation of international space station electrical performance model via on-orbit telemetry

The first U.S. power module on international space station (ISS) was activated in December 2000. Comprised of solar arrays, nickel-hydrogen (NiH/sub 2/) batteries and a direct current power management and distribution (PMAD) system, the electric power system (EPS) supplies power to housekeeping and user electrical loads. Modeling EPS performance is needed for several reasons, but primarily to assess near-term planned and off-nominal operations, and because the EPS configuration changes over the life of the ISS. The system power analysis for capability evaluation (SPACE) computer code is used to assess the ISS EPS performance. This work describes the process of validating the SPACE EPS model via ISS on-orbit telemetry. To accomplish this goal, telemetry was first used to correct assumptions and component models in SPACE. Then on-orbit data was directly input to SPACE to facilitate comparing model predictions to telemetry. It will be shown that SPACE accurately predicts on-orbit component and system performance. For example, battery state-of-charge was predicted to within 0.6 percentage points over a 0 to 100% scale, and solar array current was predicted to within a root mean square (RMS) error of 5.1 Amps out of a typical maximum of 220 Amps. First, SPACE model predictions are compared to telemetry for the ISS EPS components: solar arrays, NiH/sub 2/ batteries, and the PMAD system. Second, SPACE predictions for the overall performance of the ISS EPS are compared to telemetry and again demonstrate model accuracy.

Solar Electric Propulsion Vehicle Design Study for Cargo Transfer to Earth-Moon L1

ABSTRACTA design study for a cargo transfer vehicle using solar electric propulsion was performed for NASA’s Revolu-tionary Aerospace Systems Conceptsprogram. Targetedfor 2016, the solar electric propulsion (SEP) transfervehicle is required to deliver a propellant supply module with a mass ofapproximately 36 metric tons fromLow Earth Orbit to the first Earth-Moon libration point (LL1) within 270 days. Following an examination ofpropulsion and power technology options, a SEP transfer vehicle design was selected that incorporated large-area (~2700 m 2 ) thin film solar arrays and a clustered engine configuration of eight 50 kW gridded ionthrusters mountedonanarticulatedboom. Refinement of the SEP vehicle designwasperformediteratively toproperly estimate the required xenon propellant load for the out-bound orbit transfer. The SEP vehicle per-formance, including the xenon propellant estimation, was verified via the SNAP trajectory code. Further ef-fortsare underway to extendthissystem model to otherorbit transfer missions.INTRODUCTION

Experimental and Computational Investigations of Phase Change Thermal Energy Storage Canisters

Two sets of experimental data for cylindrical canisters with thermal energy storage applications were examined in this paper: 1) Ground Experiments and 2) Space Experiments. A 2-D computational model was developed for unsteady heat transfer (conduction and radiation) with phase-change. The radiation heat transfer employed a finite volume method. The following was found in this study: 1) Ground Experiments, the convection heat transfer is equally important to that of the radiation heat transfer; Radiation heat transfer in the liquid is found to be more significant than that in the void; Including the radiation heat transfer in the liquid resulted in lower temperatures (about 15 K) and increased the melting time (about 10 min.); Generally, most of the heat flow takes place in the radial direction. 2) Space Experiments, Radiation heat transfer in the void is found to be more significant than that in the liquid (exactly the opposite to the Ground Experiments); Accordingly, the location and size of the void affects the performance considerably; Including the radiation heat transfer in the void resulted in lower temperatures (about 40 K).