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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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The experimental objective was to assess the impact of knowledge reuse on spacecraft development time. A Secondary objective was to produce a comprehensive, flexible model other DNP teams could use to test their methodologies.
A review of existing information pertaining to spacecraft power processing systems and equipment was accomplished with a view towards applicability to the modularization of multi-kilowatt power processors. Power requirements for future spacecraft were determined from the NASA mission model-shuttle systems payload data study which provided the limits for modular power equipment capabilities. Three power processing systems were compared to evaluation criteria to select the system best suited for modularity. The shunt regulated direct energy transfer system was selected by this analysis for a conceptual design effort which produced equipment specifications, schematics, envelope drawings, and power module configurations.
The general need and requirements for the onboard embedded processors necessary to control and manipulate data in spacecraft systems are discussed. The current known requirements are reviewed from a user perspective, based on current practices in the spacecraft development process. The current capabilities of available processor technologies are then discussed, and these are projected to the generation of spacecraft computers currently under identified, funded development. An appraisal is provided for the current national developmental effort.
To help prevent the problem of mass reduction late in a spacecraft development program, the process of estimating spacecraft mass has been formulated in algorithmic form and incorporated in software that will allow trained users to interface with the integral design parameters data base interactively. The software will use historical data to extrapolate future values of a given time-dependent parameter and be manipulated by a dynamic graphics configuration builder. Future expansions of this work promise to automate the entire conceptual mission design process.
The subsequent dissertation represents an analysis of the impact of the Federal Budgetary Cycle upon the National Aeronautics and Space Administration (NASA) Manned Spacecraft Center (MSC) Research and Development (R&D) budget formulation process. The author's objectives may therefore be seen as the following: (1) to analyze the Federal Budgetary Cycle; (2) to analyze MSC R&D estimates and growth trends in relation to their implications on the Federal Cycle; (3) to identify relevant problems; and, (4) to-recommend solutions which display promise and feasibility. Any research involving the Federal Budgetary Cycle can well be characterized as of almost infinite scope and enormous complexity. For such reasons one must meticulously delineate all operational parameters and there-afore maintain their integrity. To do otherwise is to invite intellectual dilution and hazard a paltry effort.
This presentation describes the process of developing a spacecraft display system and integrated procedures.
Research work on ambient- and pasteurization-temperature reverse osmosis processes for wash water recovery in a spacecraft environment is reviewed, and the advantages and drawbacks of each are noted. A key requirement in each case is to provide a membrane of appropriate stability and semipermeability. Reverse osmosis systems intended for such use must also take into account the specific limitations and requirements imposed by the small volume of water to be processed and the high water recovery desired. The incorporation of advanced high-temperature membranes into specially designed modules is discussed.
Technology is being developed to process signals from distributed sensors using distributed computations. These distributed sensors provide a new feedback capability for vibration control that has not been exploited. Additionally, the sensors proposed are of an optical and distributed nature and could be employed with known techniques of distributed optical computation (Fourier optics, etc.) to accomplish the control system functions of filtering and regulation in a distributed computer. This paper reviews a procedure for the analytic design of control systems for this application. For illustration, the procedure is applied to the problem of suppressing the vibrations of a simply supported beam. A simulator has been developed to study the effects of sensor and processing errors. An extensive study of the effects of these errors on estimation and regulation performance is presented.
A four component system was developed which consists of a particle size reduction mechanism, a pneumatic waste transport system, a rotating-paddle incinerator, and a catalytic afterburner to be integrated into a six-man, zero-g subsystem for processing human wastes on board spacecraft. The study included the development of different concepts or functions, the establishment of operational specifications, and a critical evaluation for each of the four components. A series of laboratory tests was run, and a baseline subsystem design was established. An operational specification was also written in preparation for detailed design and testing of this baseline subsystem.
NASA's ISS and Spacecraft Processing Directorate (UB) is charged with the performance of payload development for research originating through NASA, ISS international partners, and the National Laboratory. The Payload Development sector of the Directorate takes biological research approved for on orbit experimentation from its infancy stage and finds a way to integrate and implement that research into a payload on either a Shuttle sortie or Space Station increment. From solicitation and selection, to definition, to verification, to integration and finally to operations and analysis, Payload Development is there every step of the way. My specific work as an intern this summer has consisted of investigating data received by separate flight and ground control Advanced Biological Research Systems (ABRS) units for Advanced Plant Experiments (APEX) and Cambium research. By correlation and analysis of this data and specific logbook information I have been working to explain changes in environmental conditions on both the flight and ground control unit. I have then, compiled all of that information into a form that can be presentable to the Principal Investigator (PI). This compilation allows that PI scientist to support their findings and add merit to their research. It also allows us, as the Payload Developers, to further inspect the ABRS unit and its performance
Over the past fifteen years, Goddard Space Flight Center has developed several successful science missions in-house: the Wilkinson Microwave Anisotropy Probe (WMAP), the Imager for Magnetopause-to-Aurora Global Exploration (IMAGE), the Earth Observing 1 (EO-1) [1], and the Space Technology 5 (ST-5)[2] missions, several Small Explorers, and several balloon missions. Currently in development are the Solar Dynamics Observatory (SDO) [3] and the Lunar Reconnaissance Orbiter (LRO)[4]. What is not well known is that these missions have been supported during spacecraft and/or instrument integration and test, flight software development, and mission operations by two in house satellite Telemetry and Command (T & C) Systems, the Integrated Test and Operations System (ITOS) and the Advanced Spacecraft Integration and System Test (ASIST). The advantages of an in-house satellite Telemetry and Command system are primarily in the flexibility of management and maintenance - the developers are considered a part of the mission team, get involved early in the development process of the spacecraft and mission operations-control center, and provide on-site, on-call support that goes beyond Help Desk and simple software fixes. On the other hand, care must be taken to ensure that the system remains generic enough for cost effective re-use from one mission to the next. The software is designed such that many features are user-configurable. Where user-configurable options were impractical, features were designed so as to be easy for the development team to modify. Adding support for a new ground message header, for example, is a one-day effort because of the software framework on which that code rests. This paper will discuss the many features of the Goddard satellite Telemetry and Command systems that have contributed to the success of the missions listed above. These features include flexible user interfaces, distributed parallel commanding and telemetry decommutation, a procedure language, the interfaces and tools needed for a high degree of automation, and instantly accessible archives of spacecraft telemetry. It will discuss some of the problems overcome during development, including secure commanding over networks or the Internet, constellation support for the three satellites that comprise the ST-5 mission, and geographically distributed telemetry end users.
This paper will present test data and discussion on the work we are conducting at JPL to address the following issues: 1) efficacy of sterilization process; 2) diffusion of hydrogen peroxide under sterilization process conditions into hard to reach places; 3) materials and components compatibility with the sterilization process and 4) development of methodology to protect sensitive components from hydrogen peroxide vapor.
Throughout the spacecraft design and development process, detailed instrument thermal models are created to simulate their on-orbit behavior and to ensure that they do not exceed any thermal limits. These detailed models, while generating highly accurate predictions, can sometimes lead to long simulation run times, especially when integrated with a spacecraft observatory model. Therefore, reduced models containing less detail are typically produced in tandem with the detailed models so that results may be more readily available, albeit less accurate. In the current study, both reduced and detailed instrument models are integrated with their associated spacecraft bus models to examine the impact of instrument model reduction on run time and accuracy. Preexisting instrument bus thermal model pairs from several projects were used to determine trends between detailed and reduced thermal models; namely, the Mirror Optical Bench (MOB) on the Gravity and Extreme Magnetism Small Explorer (GEMS) spacecraft, Advanced Topography Laser Altimeter System (ATLAS) on the Ice, Cloud, and Elevation Satellite 2 (ICESat-2), and the Neutral Mass Spectrometer (NMS) on the Lunar Atmosphere and Dust Environment Explorer (LADEE). Hot and cold cases were run for each model to capture the behavior of the models at both thermal extremes. It was found that, though decreasing the number of nodes from a detailed to reduced model brought about a reduction in the run-time, a large time savings was not observed, nor was it a linear relationship between the percentage of nodes reduced and time saved. However, significant losses in accuracy were observed with greater model reduction. It was found that while reduced models are useful in decreasing run time, there exists a threshold of reduction where, once exceeded, the loss in accuracy outweighs the benefit from reduced model runtime.
The energy spectra and angular distributions of electrons observed by Pioneer 11 as a function of radial distance in the inner magnetosphere of Saturn are reanalyzed and phase space densities are then calculated. The radial dependence of phase space density requires a distributed loss process. The loss is greatest in the region of the E ring (5.5 less than L less than 8.5) and is attributed to collisions with the ring particles in agreement with earlier work by Van Allen et al. (1980b). Quantitative analysis yields the following properties of the E ring: the particle radii are in the range of 4 x 10(exp -5) to 3.2 x 10(exp -4) cm and the thickness of the ring is approximately 3 R(sub s). Between the inner edge of the E ring (5.5 R(sub s)) and the outer edge of the A ring (2.3 R(sub s)) there are more energetic electrons than can be supplied by radial diffusion from an external source. Detailed calculations show that a cosmic ray albedo neutron decay (CRAND) source in the A and B rings is a plausible source for this excess. The radial diffusion coefficient required to explain the E ring absorption and CRAND source for electrons is 1 x 10(exp -12) greater than D(sub 0) greater than 3 x 10(exp -12) R(sub s)(exp 2)/s, assuming that D(sub LL) = D(sub 0)L(exp 3). As part of the reanalysis program, a method for the deconvolution of pitch angle distributions observed by simple detectors on a rotating spacecraft is developed. This process removes the instrumental response and rotational smear due to finite sampling periods and yields true angular distributions.
The energy spectra and angular distributions of electrons observed by Pioneer 11 as a function of radial distance in the inner magnetosphere of Saturn are reanalyzed and phase space densities are then calculated. The radial dependence of phase space density requires a distributed loss process. The loss is greatest in the region of the E ring (5.5 less than L less than 8.5) and is attributed to collisions with the ring particles in agreement with earlier work by Van Allen et al. (1980). Quantitative analysis yields the following properties of the E ring: the particle radii are in the range of 4 x 10(exp -5) to 3.2 x 10(exp -4) cm and the thickness of the ring is approximately 3 R(sub s). Between the inner edge of the E ring (5.5 R(sub s)) and the outer edge of the A ring (2.3 R(sub s)) there are more energetic electrons than can be supplied by radial diffusion from an external source. Detailed calculations show that a cosmic ray albedo neutron decay (CRAND) source in the A and B rings is a plausible source for this excess. The radial diffusion coefficient required to explain the E ring absorption and CRAND source for electrons is 1 x 10(exp -12) greater than D(sub 0) greater than 3 x 10(exp -12) R(exp 2, sub s)/s, assuming that D(sub LL) = D(sub 0)L(exp 3). As part of the reanalysis program, a method for the deconvolution of pitch angle distributions observed by simple detectors on a rotating spacecraft is developed. This process removes the instrumental response and rotational smear due to finite sampling periods and yields true angular distributions.
The development of a framework and structure for shuttle era unmanned spacecraft projects and the development of a commonality evaluation model is documented. The methodology developed for model utilization in performing cost trades and comparative evaluations for commonality studies is discussed. The model framework consists of categories of activities associated with the spacecraft system's development process. The model structure describes the physical elements to be treated as separate identifiable entities. Cost estimating relationships for subsystem and program-level components were calculated.
At Lockheed Missiles & Space Company, the IRIDIUM commercial space program is dramatically revolutionizing spacecraft development and manufacturing processes to reduce cost while maintaining quality and reliability. This report includes the following sections: an overview of the IRIDIUM system, the Lockheed IRIDIUM project and challenges; cycle-time reduction through production reorganization; and design for manufacturing and quality.