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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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This paper summarizes NASA/JPL progress on sub-microradian pointing system design. Sub-microradian pointing has been found to be critical for the deep space optical communications from earlier studies.
In recognition of a pressing need of the 1980s to optimize the two-way flow of information between a ground-based user and a remote-space-based sensor, an end-to-end approach to the design of information systems has been adopted at the JPL. This paper reviews End-to-End Information System (EEIS) activity at the JPL, with attention given to the scope of the EEIS transfer function, and functional and physical elements of the EEIS. The relationship between the EEIS and the NASA End-to-End Data System program is discussed.
A definition for human factors or ergonomics and its industrial and domestic application is presented. Human factors engineering, which discovers and applies information about human abilities, limitations, and other characteristics to the design of tools, machines, systems, tasks, jobs, and environments for safe, comfortable, and effective human use, is outlined. The origins of human factors and ergonomics, the philosophy of human factors, goals and objectives, systems development and design, are reviewed.
Considerations fundamental to the design of optical systems for interplanetary platforms are discussed with regard to the degree that they require departures from ground-based optical technology. Spectral availability, radiation and thermal cycles, microgravity, and space transportation constraints are considered, and new optical technologies are reviewed for their potential impact on the planetary missions planned for the rest of this century.
The purpose of this paper is to help the GAS experimenter to design a battery system that meets mission success requirements while at the same time reducing the hazards associated with the battery system. Lead-acid, silver-zinc and alkaline chemistry batteries will be discussed. Lithium batteries will be briefly discussed with emphasis on back-up power supply capabilities. The hazards associated with different battery configurations will be discussed along with the controls necessary to make the battery system two-fault tolerant.
There are a number of interesting mobile satellite applications that require the transmission of short packets of data. In the design of such systems one of the challenges is often to minimize the transmitted power to reduce cost, which implies using power efficient low bit-rate modulation and coding methods. PSK systems can be very power efficient, but at low bit-rates the carrier recovery circuits can be sensitive to oscillator phase noise. In this paper we address the problem of determining the lowest bit-rate that can be supported using PSK for a given level of system phase noise. The classical formulas are reviewed, and a method is derived to calculate the minimum C/N(sub 0) required to recover the carrier for CW, BPSK and QPSK signals for a given phase noise level.
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A system level analysis, inclusive of mass, is carried out for a cryogenic hydrogen fueled hybrid solid oxide fuel cell and bottoming gas turbine (SOFC/GT) power system. The system is designed to provide primary or secondary electrical power for an unmanned aerial vehicle (UAV) over a high altitude, long endurance mission. The net power level and altitude are parametrically varied to examine their effect on total system mass. Some of the more important technology parameters, including turbomachinery efficiencies and the SOFC area specific resistance, are also studied for their effect on total system mass. Finally, two different solid oxide cell designs are compared to show the importance of the individual solid oxide cell design on the overall system. We show that for long mission durations of 10 days or more, the fuel mass savings resulting from the high efficiency of a SOFC/GT system more than offset the larger powerplant mass resulting from the low specific power of the SOFC/GT system. These missions therefore favor high efficiency, low power density systems, characteristics typical of fuel cell systems in general.
To provide the transportation of lunar base elements to the moon, large high-energy propulsion systems will be required. Advanced propulsion systems for lunar missions can provide significant launch mass reductions and payload increases. These mass reductions and added payload masses can be translated into significant launch cost savings for the lunar base missions. The masses in low Earth orbit (LEO) were compared for several propulsion systems: nitrogen tetroxide/monomethyl hydrazine (NTO/MMH), oxygen/methane (O2/CH4), oxygen/hydrogen (O2/H2), and metallized O2/H2/Al propellants. Also, the payload mass increases enabled with O2/H2 and O2/H2/Al systems were addressed. In addition, many system design issues involving the engine thrust levels, engine commonality between the transfer vehicle and the excursion vehicle, and the number of launches to place the lunar mission vehicles into LEO will be discussed. Analyses of small lunar missions launched from a single STS-C flight are also presented.
The NASA Safety Standard, which establishes a uniform process for hydrogen system design, materials selection, operation, storage, and transportation, is presented. The guidelines include suggestions for safely storing, handling, and using hydrogen in gaseous (GH2), liquid (LH2), or slush (SLH2) form whether used as a propellant or non-propellant. The handbook contains 9 chapters detailing properties and hazards, facility design, design of components, materials compatibility, detection, and transportation. Chapter 10 serves as a reference and the appendices contained therein include: assessment examples; scaling laws, explosions, blast effects, and fragmentation; codes, standards, and NASA directives; and relief devices along with a list of tables and figures, abbreviations, a glossary and an index for ease of use. The intent of the handbook is to provide enough information that it can be used alone, but at the same time, reference data sources that can provide much more detail if required.
This report describes the results of research activities associated with the development of rapid design systems for aerospace structures in support of the Intelligent Synthesis Environment (ISE). The specific subsystems investigated were the interface between model assembly and analysis; and, the high performance NASA GPS equation solver software system in the Windows NT environment on low cost high-performance PCs.
After reviewing third-order aberration theory, three examples are given of the design of lenses by the concentric-aplanatic method. The performance of any optical system designed with this method can be improved by releasing all the constraints one by one. After correcting for third-order aberrations, the error function can be inspected to determine if the system is dominated by third- or higher-order aberrations.
This paper describes the design of a wideband S-band communications link from the Space Shuttle Orbiter to a ground-based terminal. This link will be used for transmission of television, digital data (various rates), or analog data, and is of interest because of the presence of bandwidth, power, and hardware constraints. The procedure followed to arrive at an overall signal and system design is outlined, and system performance margins are presented for various operational configurations. Finally, the growth potential of this link for digital data transmission is discussed.
This paper describes the attitude ground system (AGS) design to be used for support of the Magnetospheric MultiScale (MMS) mission. The AGS exists as one component of the mission operations control center. It has responsibility for validating the onboard attitude and accelerometer bias estimates, calibrating the attitude sensors and the spacecraft inertia tensor, and generating a definitive attitude history for use by the science teams. NASA's Goddard Space Flight Center (GSFC) in Greenbelt, Maryland is responsible for developing the MMS spacecraft, for the overall management of the MMS mission, and for mission operations. MMS is scheduled for launch in 2014 for a planned two-year mission. The MMS mission consists of four identical spacecraft flying in a tetrahedral formation in an eccentric Earth orbit. The relatively tight formation, ranging from 10 to 400 km, will provide coordinated observations giving insight into small-scale magnetic field reconnection processes. By varying the size of the tetrahedron and the orbital semi-major axis and eccentricity, and making use of the changing solar phase, this geometry allows for the study of both bow shock and magnetotail plasma physics, including acceleration, reconnection, and turbulence. The mission divides into two phases for science; these phases will have orbit dimensions of 1.2 x 12 Earth radii in the first phase and 1.2x25 Earth radii in the second in order to study the dayside magnetopause and the nightside magnetotail, respectively. The orbital periods are roughly one day and three days for the two mission phases. Each of the four MMS spacecraft will be spin stabilized at 3 revolutions per minute (rpm), with the spin axis oriented near the ecliptic north pole but tipped approximately 2.5 deg towards the Sun line. The main body of each spacecraft will be an eight-sided platform with diameter of 3.4 m and height of 1.2 m. Several booms are attached to this central core: two axial booms of 14.9 m length, two radial magnetometer booms of 5 m length, and four radial wire booms of 60 m length. Attitude and orbit control will use a set of axial and radial thrusters. A four-head star tracker and a slit-type digital Sun sensor (DSS) provide input for attitude determination. In addition, an accelerometer will be used for closed-loop orbit maneuver control. The primary AGS product will be a daily definitive attitude history. Due to power limitations, the star tracker and accelerometer data will not be available at all times. However, tracker data from at least 10 percent of each orbit and continuous DSS data will be provided. An extended Kalman filter (EKF) will be used to estimate the three-axis attitude (i.e., spin axis orientation and spin phase) and rotation rate for all times when the tracker data is valid. For other times, the attitude is generated by assuming a constant angular momentum vector in the inertial frame. The DSS sun pulse will provide a timing signal to maintain an accurate spin phase. There will be times when the Sun is occulted and DSS data is not available. If this occurs at the start or end of a definitive attitude product, then the spin phase will be extrapolated using the mean rate determined by the EKF.
This paper describes the attitude ground system (AGS) design to be used for support of the Magnetospheric MultiScale (MMS) mission. The AGS exists as one component of the mission operations control center. It has responsibility for validating the onboard attitude and accelerometer bias estimates, calibrating the attitude sensors and the spacecraft inertia tensor, and generating a definitive attitude history for use by the science teams. NASA's Goddard Space Flight Center (GSFC) in Greenbelt, Maryland is responsible for developing the MMS spacecraft, for the overall management of the MMS mission, and for mission operations. MMS is scheduled for launch in 2014 for a planned two-year mission. The MMS mission consists of four identical spacecraft flying in a tetrahedral formation in an eccentric Earth orbit. The relatively tight formation, ranging from 10 to 400 km, will provide coordinated observations giving insight into small-scale magnetic field reconnection processes. By varying the size of the tetrahedron and the orbital semi-major axis and eccentricity, and making use of the changing solar phase, this geometry allows for the study of both bow shock and magnetotail plasma physics, including acceleration, reconnection, and turbulence. The mission divides into two phases for science; these phases will have orbit dimensions of l.2xl2 Earth radii in the first phase and l.2x25 Earth radii in the second in order to study the dayside magnetopause and the nightside magnetotail, respectively. The orbital periods are roughly one day and three days for the two mission phases. Each of the four MMS spacecraft will be spin stabilized at 3 revolutions per minute (rpm), with the spin axis oriented near the ecliptic north pole but tipped approximately 2.5 deg towards the Sun line. The main body of each spacecraft will be an eight-sided platform with diameter of 3.4 m and height of 1.2 m. Several booms are attached to this central core: two axial booms of 14.9 m length, two radial magnetometer booms of 5 m length, and four radial -wire booms of 60 m length. Attitude and orbit control will use a set of axial and radial thrusters. A four-head star tracker and a slit-type digital Sun sensor (DSS) provide input for attitude determination. In addition, an accelerometer will be· used for closed-loop orbit maneuver control. The primary AGS product will be a daily definitive attitude history. Due to power limitations; the star tracker and accelerometer data will not be available at all times. However, tracker data from at least 10 percent of each orbit and continuous DSS data will be provided. An extended Kalman filter (EKF) will be used to estimate the three-axis attitude (i.e., spin axis orientation and spin phase) and rotation rate for all times when the tracker data is valid. For other times, the attitude is generated by assuming a constant angular momentum vector in the inertial frame. The DSS sun pulse will provide a timing signal to maintain an accurate spin phase. There will be times when the Sun is occulted and DSS data is not available. If this occurs at the start or end of a definitive attitude product, then the spin phase will be extrapolated using the mean rate determined by the EKF.
This paper describes the system architecture, design and implementation approach for the Mars Pathfinder spacecraft scheduled to land on the surface of Mars on July 4, 1997. Mars Pathfinder is one of the new series of small, challenging missions doing significant science/engineering on a fast schedule and cost capped budget.
This paper illustrates how traditional analysis and design tools can be successfully applied, with insights derived from recent research on control-structure interaction, to preliminary control system design for a large space structure such as the Space Station Freedom. The control system regulates attitude during construction, normal on-orbit operation, reboost, and docking using control-moment gyros and thrusters. The tools include finite-element modeling, model reduction based on modal truncation, single-input/single-output Bode plot design, position and rate feedback, roll-off and phase filter, and simulation. The resulting system is characterized by simple architecture and moderate bandwidth. The large configuration changes during the construction period are handled by gain scheduling.
The High Performance Antiproton Trap (HiPAT), a system designed to hold up to l0(exp 12) charge particles with a storage half-life of approximately 18 days, is a tool to support basic antimatter research. NASA's interest stems from the energy density represented by the annihilation of matter with antimatter, 10(exp 2)MJ/g. The HiPAT is configured with a Penning-Malmberg style electromagnetic confinement region with field strengths up to 4 Tesla, and 20kV. To date a series of normal matter experiments, using positive and negative ions, have been performed evaluating the designs performance prior to operations with antiprotons. The primary methods of detecting and monitoring stored normal matter ions and antiprotons within the trap includes a destructive extraction technique that makes use of a micro channel plate (MCP) device and a non-destractive radio frequency scheme tuned to key particle frequencies. However, an independent means of detecting stored antiprotons is possible by making use of the actual annihilation products as a unique indicator. The immediate yield of the annihilation event includes photons and pie mesons, emanating spherically from the point of annihilation. To "count" these events, a hardware system of scintillators, discriminators, coincident meters and multi channel scalars (MCS) have been configured to surround much of the HiPAT. Signal coincidence with voting logic is an essential part of this system, necessary to weed out the single cosmic ray events from the multi-particle annihilation shower. This system can be operated in a variety of modes accommodating various conditions. The first is a low-speed sampling interval that monitors the background loss or "evaporation" rate of antiprotons held in the trap during long storage periods; provides an independent method of validating particle lifetimes. The second is a high-speed sample rate accumulating information on a microseconds time-scale; useful when trapped antiparticles are extracted against a target, providing an indication of quantity. This paper details the layout of this system, setup of the hardware components around HiPAT, and applicable checkouts using normal matter radioactive sources.