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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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At least 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Spacecraft systems design trade-offs for the Earth Resources Technology Satellite.

The Earth Resources Technology Satellite Program's use of flight proven hardware in the design of a satellite for earth sensor payload support and data handling is discussed. The use of an existing satellite as the building block around which additional support systems such as the orbit adjust system, the redundant wideband telemetry systems, the second regulated power system, and the quad redundant command system is analyzed. System performance seen in orbit vs design objectives are discussed to point up the success of the design approach chosen. Also discussed are the schedule and cost benefits derived from the use of previously developed hardware with additional subsystems as required to meet program requirements.

Branchflower, G. A.↗
NASA NTRS1975

Reliability considerations in long-life outer planet spacecraft system design

A Mariner Jupiter/Saturn mission has been planned for 1977. System reliability questions are discussed, taking into account the actual and design lifetime, causes of mission termination, in-flight failures and their consequences for the mission, and the use of redundancy to avoid failures. The design process employed optimizes the use of proven subsystem and system designs and then makes the necessary improvements to increase the lifetime as required.

Casani, E. K.↗
NASA NTRS1978

Investigation of high voltage spacecraft system interactions with plasma environments

An experimental investigation was undertaken for insulator and conductor test surfaces biased up to + or - 1kV in a simulated low earth orbit charged particle environment. It was found that these interactions are controlled by the insulator surfaces surrounding the biased conductors. For positive applied voltages the electron current collection can be enhanced by the insulators. For negative applied voltages the insulator surface confines the voltage to the conductor region. Understanding these interactions and the technology to control their impact on system operation is essential to the design of solar cell arrays for ion drive propulsion applications that use direct drive power processing.

Stevens, N. J.↗
NASA NTRS1979

Stability and control of future spacecraft systems

A brief review is presented of some of the problem areas associated with large space structures and some of the approaches currently being taken to find solutions are discussed. Consideration is given to such areas as control system analysis, design and implementation, and those aspects of structural modeling related to control system design.

Szirmay, S. Z.↗
NASA NTRS1983

The Galileo spacecraft system design

The Galileo Orbiter differs markedly from its Mariner, Viking and Voyager predecessors in many aspects of its design, as dictated by its unique mission characteristics. Novel design features include a dual spin configuration, long life bipropellant propulsion, radioisotope thermoelectric generators, and a star scanner. Other major changes are related to Space Shuttle Orbiter compatibility. The selection of a distributed data system employing a high speed data bus represents a major advance over Voyager systems. The command, control and data handling functions for the Orbiter are merged into a single software-controlled subsystem. The paramount importance of software development in the design of Galileo subsystems is noted, in view of the requirement for total engineering and science data acquisition process reprogrammability.

Jones, C. P.↗
NASA NTRS1986

Spacecraft Systems Working Group report

Issues addressed include: definition of user/commercial/government needs by function; criteria for prioritization of needs; overall criteria for technology assessment; system configuration drivers (key trade studies); space infrastructure interface; and cost drivers (pros and cons of standardization, manufacturing, test, serviceability, and supportability).

Keigler, John E.↗
NASA NTRS1990

How much redundancy: Some cost considerations, including examples for spacecraft systems

How much redundancy should be built into a subsystem such as a space power subsystem. How does a reliability or design engineer choose between a power subsystem with 0.990 reliability and a more costly subsystem with 0.995 reliability. How does the engineer designing a power subsystem for a satellite decide between one power subsystem and a more reliable but heavier power subsystem. High reliability is not necessarily an end in itself. High reliability may be desirable in order to reduce the statistically expected loss due to a subsystem failure. However, this may not be the wisest use of funds since the expected loss due to subsystem failure is not the only cost involved. The subsystem itself may be very costly. The cost of the subsystem or the expected loss due to subsystem failure may not be considered separately. Therefore, the total of the two costs is minimized, i.e., the total of the cost of the subsystem plus the expected loss due to subsystem failure. A specific type of redundant system is considered, called a k-out-of-n: G subsystem. Such a subsystem has n modules, of which k are required to be good for the subsystem to be good. Five models are discussed which can be applied in the design of a power subsystem to select the unique redundancy method which will minimize the total of the cost of the power subsystem plus the expected loss due to the power subsystem failure. A BASIC computer program is available.

Suich, Ronald C.↗
NASA NTRS1993

Spacecraft systems engineering: An introduction to the process at GSFC

The main objective in systems engineering is to devise a coherent total system design capable of achieving the stated requirements. Requirements should be rigid. However, they should be continuously challenged, rechallenged and/or validated. The systems engineer must specify every requirement in order to design, document, implement and conduct the mission. Each and every requirement must be logically considered, traceable and evaluated through various analysis and trade studies in a total systems design. Margins must be determined to be realistic as well as adequate. The systems engineer must also continuously close the loop and verify system performance against the requirements. The fundamental role of the systems engineer, however, is to engineer, not manage. Yet, in large, complex missions, where more than one systems engineer is required, someone needs to manage the systems engineers, and we call them 'systems managers.' Systems engineering management is an overview function which plans, guides, monitors and controls the technical execution of a project as implemented by the systems engineers. As the project moves on through Phases A and B into Phase C/D, the systems engineering tasks become a small portion of the total effort. The systems management role increases since discipline subsystem engineers are conducting analyses and reviewing test data for final review and acceptance by the systems managers.

Fragomeni, Tony↗
NASA NTRS1992

Microencapsulated PCM slurries for heat transfer and energy storage in spacecraft systems

The technical feasibility for providing significantly enhanced heat transport and storage as well as improved thermal control has been investigated during several Small Business Innovative Research (SBIR) programs for NASA, the United States Air Force (USAF), and the Strategic Defense Initiative Organization (SDIO) using microencapsulated phase change materials (PCMs) in both aqueous and nonaqueous two-component slurries. In the program for SDIO, novel two-component coolant fluids were prepared and successfully tested at both low (300 K) and intermediate temperatures (460 to 700 K). The two-component fluid slurries of microencapsulated PCMs included organic particles in aqueous and nonaqueous liquids, as well as microencapsulated metals that potentially could be carried by liquid metals or used as powdered heat sinks. Simulation and experimental studies showed that such active cooling systems could be designed and operated with enhancements of heat capacity that exceeded 10 times or 1000 percent that for the base fluid along with significant enhancement in the fluid's heat capacity. Furthermore, this enhancement provided essentially isothermal conditions throughout the pumped primary coolant fluid loop. The results suggest that together with much higher fluid thermal capacity, greater uniformity of temperature is achievable with such fluids, and that significant reductions in pumping power, system size, and system mass are also possible.

Colvin, David P.↗
NASA NTRS1993

Low power arcjet system spacecraft impacts

Potential plume contamination of spacecraft surfaces was investigated by positioning spacecraft material samples relative to an arcjet thruster. Samples in the simulated solar array region were exposed to the cold gas arcjet plume for 40 hrs to address concerns about contamination by backstreaming diffusion pump oil. Except for one sample, no significant changes were measured in absorptance and emittance within experimental error. Concerns about surface property degradation due to electrostatic discharges led to the investigation of the discharge phenomenon of charged samples during arcjet ignition. Short duration exposure of charged samples demonstrated that potential differences are consistently and completely eliminated within the first second of exposure to a weakly ionized plume. The spark discharge mechanism was not the discharge phenomenon. The results suggest that the arcjet could act as a charge control device on spacecraft.

Pencil, Eric J.↗
NASA NTRS1993

Low power arcjet system spacecraft impacts

Application of electrothermal arcjets on communications satellites requires assessment of integration concerns identified by the user community. Perceived risks include plume contamination of spacecraft materials, induced arcing or electrostatic discharges between differentially charged spacecraft surfaces, and conducted and radiated electromagnetic interference (EMI) for both steady state and transient conditions. A Space Act agreement between Martin Marietta Astro Space, the Rocket Research Company, and NASA's Lewis Research Center was established to experimentally examine these issues. Spacecraft materials were exposed to an arcjet plume for 40 hours, representing 40 weeks of actual spacecraft life, and contamination was characterized by changes in surface properties. With the exception of the change in emittance of one sample, all measurable changes in surface properties resulted in acceptable end of life characteristics. Charged spacecraft samples were benignly and consistently reduced to ground potential during exposure to the powered arcjet plume, suggesting that the arcjet could act as a charge control device on spacecraft. Steady state EMI signatures obtained using two different power processing units were similar to emissions measured in a previous test. Emissions measured in UHF, S, C, Ku and Ka bands obtained a null result which verified previous work in the UHF, S, and C bands. Characteristics of conducted and radiated transient emissions appear within standard spacecraft susceptibility criteria.

Pencil, Eric J.↗