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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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Hot-water system requires no external power except solar energy. System is completely self-controlling. It includes solar-powered pump, solar-thermally and hydrothermally operated valves, and storage tank filled with open-celled foam, to maintain thermal stratification in stored water.
Lightweight, multikilowatt solar power arrays in conjunction with electric propulsion offer potential improvements to space exploration, extending the usefulness of existing launch vehicles to higher-energy missions. Characteristics of solar-powered electric propulsion missions are outlined, and preliminary performance estimates are shown. Spacecraft system engineering is discussed with respect to parametric trade-offs in power and propulsion system design. Relationships between mission performance and propulsion system performance are illustrated. The present state of the art of electric propulsion systems is reviewed and related to the mission requirements identified earlier. The propulsion system design and test requirements for a mission spacecraft are identified and discussed. Although only ion engine systems are currently available, certain plasma propulsion systems offer some advantages in over-all system design. These are identified, and goals are set for plasma-thrustor systems to make them competitive with ion-engine systems for mission applications.
Solar electric power system for ion propelled Jupiter flyby mission, noting photovoltaic solar array, battery energy storage and power conditioning
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Concepts developed for both LEO and GEO construction of photovoltaic and thermal engine satellites are analyzed. Topics discussed include: satellite construction; crew scheduling; crew jobs and organizations; operator productivity rating; constructability rating; transportation systems for cargo launch, refueling operations, personnel transport, and orbit transfer; collision analysis, cost analysis, and radiation evironment and effects.
The best estimates of space transportation requirements for cargo launch vehicles, personnel launch carriers, high thrust orbit transfer, and electric orbit transfer systems are discussed, along with the rationale for each.
Both LEO transportation (earth to low earth orbit) and GEO transportation (low earth orbit to geosynchronous orbit) segments were addressed. LEO options include both a 2 stage winged space freighter vehicle and a 2 stage ballistic recoverable vehicle. Both incorporate LO(2)/RP-1/LH(2) engines on the booster and standard SSME's on the upper stage. The orbit transfer vehicle options included chemical for geosynchronous satellite assembly and self powered electric propulsion for low earth orbit satellite assembly. An exhaust products analysis was conducted for the earth to LEO vehicle since atmospheric pollution could be a concern.
Practical designs for power transmission were developed to meet requirements and constraints. Microwave link error was analyzed to confirm attainability of acceptable link efficiency. Silicon photovoltaic was determined to be the best overall choice for energy conversion, with a potassium Rankine cycle as the backup choice. Space transportation operations provide low cost because of traffic level, and the payload volume is the launch vehicle design driver. The power cost is 4 to 5 /kwh, which will be competitive with fossil fuel sources by the year 2000.
A slotted waveguide planar array was established as the baseline design for the spaceborne transmitter antenna. Key aspects of efficient energy conversion at both ends of the power transfer link were analyzed and optimized alternate approaches in the areas of antenna and tube design are discussed. An integrated design concept was developed which meets design requirements, observes structural and thermal constraints, exhibits good performance and was developed in adequate depth to permit cost estimating at the subsystem/component level.
Construction and transportation systems and operations are described for the following combinations: (1) silicon photovoltaic CR=1 satellite constructed primarily in low earth orbit (LEO); (2) silicon photovoltaic CR=1 satellite constructed in geosynchronous earth orbit (GEO); (3) Rankine thermal engine satellite constructed primarily in LEO; and (4) Rankine thermal engine satellite constructed in GEO.
A preliminary maintenance analysis was made and system maintenance costs were included. The far sidelobes for the SPS transmitters were studied and demonstration options for the SPS were evaluated.
Various approaches to the problem of generating, maintaining and distributing a coherent, reference phase signal over a large area are suggested, mathematically modeled and analyzed with respect to their ability to minimize: phase build-up, beam diffusion and beam steering phase jitter, cable length, and maximize power transfer efficiency. In addition, phase control configurations are suggested which alleviate the need for layout symmetry.
The goal of the Project is to support the industrial development of PFDR technology that will provide favorable life-cycle costs per unit of electrical or thermal energy produced. The technology will be made available in the early 1980s for applications project experiments. PFDR systems utilize concentrator dishes to furnish energy to their own individual receivers and power conversion subsystems. Initial effort is with steam Rankine and gas Brayton cycles. Periodic assessments will be made to confirm or change the cycles initially selected. Subsystems will be designed, fabricated and tested together in modules as appropriate. This paper describes PFDR systems briefly, outlines the project goals and organization, discusses the plans and current status of the project, and lists the benefits of PFDR technology concepts.
The results of the operations analyses are reported. Some of these analyses examined operations aspects of space vehicle in-space maintenance. Many of the analyses explored in great depth operations concerned the LEO Base cargo handling operations. Personnel transportation operations and cargo packaging were also analyzed. These operations analyses were performed to define the operational requirements for all of the SPS system elements so that equipment and facilities could be synthesized, and to make estimates of the manpower requirements. An overall, integrated, end-to-end description of the SPS operations is presented. The detailed operations analyses, upon which this integrated description was based, are included.
Proposed space missions will require large scale, light weight, space based structural systems. Large space structure technology (LSST) systems will have to accommodate (among others): ocean data systems; electronic mail systems; large multibeam antenna systems; and, space based solar power systems. The structures are to be delivered into orbit by the space shuttle. Because of their inherent size, modelling techniques and scaling algorithms must be developed so that system performance can be predicted accurately prior to launch and assembly. When the size and weight-to-area ratio of proposed LSST systems dictate that the entire system be considered flexible, there are two basic modeling methods which can be used. The first is a continuum approach, a mathematical formulation for predicting the motion of a general orbiting flexible body, in which elastic deformations are considered small compared with characteristic body dimensions. This approach is based on an a priori knowledge of the frequencies and shape functions of all modes included within the system model. Alternatively, finite element techniques can be used to model the entire structure as a system of lumped masses connected by a series of (restoring) springs and possibly dampers. In addition, a computational algorithm was developed to evaluate the coefficients of the various coupling terms in the equations of motion as applied to the finite element model of the Hoop/Column.
Proposed space missions will require large scale, light weight, space based structural systems. Large space structure technology (LSST) systems will have to accommodate (among others): ocean data systems; electronic mail systems; large multibeam antenna systems; and, space based solar power systems. The structures are to be delivered into orbit by the Space Shuttle. Because of their inherent size, modelling techniques and scaling algorithms must be developed so that system performance can be predicted accurately prior to launch and assembly. When the size and weight-to-area ratio of proposed LSST systems dictate that the entire system be considered flexible, there are two basic modelling methods which can be used. The first is a continuum approach, a mathematical formulation for predicting the motion of a general orbiting flexible body, in which elastic deformations are considered small compared with characteristic body dimensions. This approach is based on an a priori knowledge of the frequencies and shape functions of all modes included within the system model. Alternatively, finite element techniques can be used to model the entire structure as a system of lumped masses connected by a series of (restoring) springs and possibly dampers. In addition, a computational algorithm was developed to evaluate the coefficients of the various coupling terms in the equations of motion as applied to the finite element model of the Hoop/Column.
Mass, area, and station-keeping propellant needs have been estimated for a typical system. And, although important criteria such as cost, Shuttle packaging, and erection/deployment schemes were not considered, the documented trends should aid in many of the design choices to be made. Effects on system characteristics were examined for: three heat storage salts with melting temperatures from 743 to 1121 K; parabolic and Cassegrainian mirrors; module power levels of 20 and 40 kW; and, alternate pumped-loop, tube-and-fin radiator configurations, with and without micrometeoroid armoring.
A computer code to model the steady-state performance of a monogroove heat pipe for the NASA Space Station is presented, including the effects on heat pipe performance of a screen in the evaporator section which deals with transient surges in the heat input. Errors in a previous code have been corrected, and the new code adds additional loss terms in order to model several different working fluids. Good agreement with existing performance curves is obtained. From a preliminary evaluation of several of the radiator design parameters it is found that an optimum fin width could be achieved but that structural considerations limit the thickness of the fin to a value above optimum.