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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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An overview of long-range options for in-space laser power transmission is presented. The focus is on the new technology and research status of solar-pumped lasers and their solar concentration needs. The laser options include gas photodissociation lasers, optically-pumped solid-state lasers, and blackbody-pumped transfer lasers. The paper concludes with a summary of current research thrusts.
An expert system called EXADS has been developed to aid users of the Automated Design Synthesis (ADS) general purpose optimization program. ADS has approximately 100 combinations of strategy, optimizer, and one-dimensional search options from which to choose. It is difficult for a nonexpert to make this choice. This expert system aids the user in choosing the best combination of options based on the users knowledge of the problem and the expert knowledge stored in the knowledge base. The knowledge base is divided into three categories; constrained problems, unconstrained problems, and constrained problems being treated as unconstrained problems. The inference engine and rules are written in LISP, contains about 200 rules, and executes on DEC-VAX (with Franz-LISP) and IBM PC (with IQ-LISP) computers.
Results are given of a study to explore options for the development of a Controlled Ecological Life Support System (CELSS) for a future Space Station. In addition, study results will benefit the design of other facilities such as the Life Sciences Research Facility, a ground-based CELSS demonstrator, and will be useful in planning longer range missions such as a lunar base or manned Mars mission. The objectives were to develop weight and cost estimates for one CELSS module selected from a set of preliminary plant growth unit (PGU) design options. Eleven Space Station CELSS module conceptual PGU designs were reviewed, components and subsystems identified and a sensitivity analysis performed. Areas where insufficient data is available were identified and divided into the categories of biological research, engineering research, and technology development. Topics which receive significant attention are lighting systems for the PGU, the use of automation within the CELSS system, and electric power requirements. Other areas examined include plant harvesting and processing, crop mix analysis, air circulation and atmosphere contaminant flow subsystems, thermal control considerations, utility routing including accessibility and maintenance, and nutrient subsystem design.
Mars Mission profile options and mission requirements data are presented for Earth-Mars opposition and conjunction class round-trip flyby and stopover mission opportunities. The opposition class flyby and sprint mission uses direct transfer trajectories to and on return from Mars. The opposition class stopover mission employs the gravitational field of Venus to accelerate the space vehicle on either the outbound or inbound leg in order to reduce the propulsion requirement associated with the opposition class mission. The conjunction class mission minimizes propulsion requirements by optimizing the stopover time at Mars. Representative interplanetary space vehicle systems are sized to compare and show sensitivity of the initial mass required in low Earth orbit to one mission profile option and mission opportunity to another.
Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment have been evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA II 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10(7) to 2.1x10(7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA II 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10(6) to 3.6x10(6) N-s, and approximately 662 to 1072 m/s, respectively.
Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment were evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA 2 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10 (exp 7) to 2.1x10 (exp 7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA 2 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10 (exp 6) to 3.6x10 (exp 6) N-s, and approximately 662 to 1072 m/s, respectively.
The paper presents launch vehicle and upper-stage options for application to lunar and interplanetary microspacecraft missions. Particular attention is given to the capabilities of Piggyback, small launch vehicles, and large launch vehicles. It is noted that Piggyback options on the Shuttle and expendable launch vehicles enable near-term earth-orbital missions and the potential for lunar and planetary missions if an electric-propulsion upper-stage is developed. Launch systems like the Space Shuttle could be used to launch large members of microspacecraft in 'constellation deployment' and 'shotgun' class missions to a variety of solar-system targets such as the sun, asteroids, comets, the moon, Mars, and Saturn.
Mars Mission profile options and mission requirements data are presented for earth-Mars opposition and conjuction class round-trip flyby and stopover mission opportunities. The opposition class flyby and sprint mission uses direct transfer trajectories to and on return from Mars. The opposition-class stopover mission employs the gravitational field of Venus to accelerate the space vehicle on either the outbound or inbound leg in order to reduce the propulsion requirement associated with the opposition-class mission. The conjuction-class mission minimizes propulsion requirements by optimizing the stopover time at Mars. Representative interplanetary space vehicle systems are sized to compare and show sensitivity of the initial mass required in low earth orbit to one mission profile option and mission opportunity to another.
In order to meet the growing power demands of the Space Station, the electrical power system design includes an option to provide additional power capability in increments of 50 kWe of power. Each increment consists of a pair of two solar dynamic power modules (SDPMs), each of which containing a closed Brayton Cycle (CBC) thermodynamic engine. A solar dynamic radiator (SDR) enables the CBC to reject the waste heat to the surrounding space environment. This paper analyzes three alternatives to the baseline configuration of the Space Station Freedom solar dynamic radiator and discusses their merits based on Space Shuttle cargo capabilities, location with respect to the SDPM supporting structure, thermal performance, drag, concentrator shading, mass, and other issues of concenrn. Results indicating the advantages and disadvantages of each option are presented along with diagrams of the alternative configurations.
Mars Mission profile options and mission requirements data are presented for earth-Mars opposition and conjunction class round-trip flyby and stopover mission opportunities. The opposition-class flyby and sprint mission uses direct transfer trajectories to and on return from Mars. The opposition-class stopover mission employs the gravitational field of Venus to accelerate the space vehicle on either the outbound or inbound leg in order to reduce the propulsion requirement associated with the opposition-class mission. The conjunction-class mission minimizes propulsion requirements by optimizing the stopover time at Mars. Representative interplanetary space vehicle systems are sized to compare and show sensitivity of the initial mass required in low earth orbit to one mission profile option and mission opportunity to another.
A universal noiseless coding structure was developed that provides efficient performance over an extremely broad range of source entropy. This is accomplished by adaptively selecting the best of several easily implemented variable length coding algorithms. Custom VLSI coder and decoder modules capable of processing over 20 million samples per second are currently under development. The first of the code options used in this module development is shown to be equivalent to a class of Huffman code under the Humblet condition, other options are shown to be equivalent to the Huffman codes of a modified Laplacian symbol set, at specified symbol entropy values. Simulation results are obtained on actual aerial imagery, and they confirm the optimality of the scheme. On sources having Gaussian or Poisson distributions, coder performance is also projected through analysis and simulation.
This paper reviews the present International Reference Ionosphere (IRI) model of electron density in the middle ionosphere and explores two new options for future editions of IRI. The first of these options is a better description of the bottomside thickness parameters, and the second is an analytical representation from E- to F2-peak using LAY-functions. For this analytical representation a table of standard parameters and constraints for the four LAY-functions recommended for IRI have been established.
The Nuclear Thermal Rocket (NTR) is an attractive candidate propulsion system option for manned planetary missions. Its high performance capability for such missions translates into a substantial reduction in low-earth-orbit (LEO) required mass and trip times with increased operational flexibility. This study examined NTR engine options that could support near-term lunar mission operations. Expander and gas generator cycle, solid-core NERVA derivative reactor-based NTR engines were investigated. Weight, size, operational characteristics, and design features for representative NTR engine concepts are presented. The impact of using these NTR engines for a typical lunar mission scenario is also examined.
The objectives were to examine launch vehicle applications and propulsion requirements for potential future manned space transportation systems and to support planning toward the evolution of Space Shuttle Main Engine (SSME) and Space Transportation Main Engine (STME) engines beyond their current or initial launch vehicle applications. As a basis for examinations of potential future manned launch vehicle applications, we used three classes of manned space transportation concepts currently under study: Space Transportation System Evolution, Personal Launch System (PLS), and Advanced Manned Launch System (AMLS). Tasks included studies of launch vehicle applications and requirements for hydrogen-oxygen rocket engines; the development of suggestions for STME engine evolution beyond the mid-1990's; the development of suggestions for STME evolution beyond the Advanced Launch System (ALS) application; the study of booster propulsion options, including LOX-Hydrocarbon options; the analysis of the prospects and requirements for utilization of a single engine configuration over the full range of vehicle applications, including manned vehicles plus ALS and Shuttle C; and a brief review of on-going and planned LOX-Hydrogen propulsion technology activities.
AFD has completed the trade study on Standard Insertion (SI) vs Direct Insertion (DI) for STS-50. RSOC Range Safety has developed acceptable DI targets from 130 n.mi. to 150 n.mi. and the corresponding performance assessment for these targets using STS-50 data has been completed. This mission has sufficient performance capability to perform this mission as a DI to 160 n.mi. A reduced OMS load corresponding to a DI mission is required for this option. The increase in altitude over the AFP baseline (SI to 145 n.mi.) is highly desirable for this mission. The orientation on orbit for the orbiter/USML-1 payload is such that orbital decay is maximized (maximum frontal cross-sectional area with vehicle normal to velocity vector). Increasing the operational altitude reduces the amount of vernier thruster firings necessary to maintain a constant gravity gradient. The results of this trade study can also be applied to other superheavyweight missions (EDO flights) and will allow for use of the DI technique for lower orbital altitudes, thereby eliminating the SI option for due east, low altitude missions. STSOC transmittal form no. 330-330-130, which documents the technical issues and assumptions used for this trade study effort in detail, should be referenced for further information. The main reason that a DI is desired for STS-50 and other superheavyweight flights (low altitude) is that ESMC range safety has expressed reservations about SI missions in general. The concern is that the current SI design underspeed exposes Africa and Madagascar to potential ET debris impact. In the past range safety has waived the requirement that these areas be protected in the event of an engine failure. With the advent of the pre-MECO OMS dump, the viability of DI and the high casualty expectations from the ACTA press to MECO hazard study, range safety has become more reluctant to approve SI flights. It is felt that to perform an SI mission there would have to be a large decrease in design underspeed to protect these landmasses, which would result in possible gaps between a late TAL and PTA. The assumed limit on a DI has been to altitudes greater than 160 n.mi. It was assumed that for altitudes less than 160 n.mi. the nominal ET impact envelope would overlap the Gilbert Islands.
A comparison of the various cooling options for the Astromag particle physics experiment is presented. The baseline design for the cryogenic system involves using a natural circulation fountain-effect driven flow loop (Hofmann type). The present paper considers two alternative options for cooling. The first design involves a thermal strap made of a high-conductivity metal, e.g., high-purity aluminum or copper, which connects the coil to the helium reservoir. Venting helium vapor can also be used to minimize the temperature of the magnet and recover from a quench. The second design is based on an He II heat pipe concept where steady state heat transport is by counterflow. Cavitation is prevented by use of a porous plug. Forced flow He II is also available but only during extraordinary operating conditions.
Lunar transportation system (LTS) options to meet defined lunar mission requirements are described. The discussed configuration and scenarios include: ground-based direct launch, ground-based earth orbit rendezvous/dock, and space-based on-orbit assembly. The earth-to-orbit vehicle sizing, operational complexities, and technology phasing of each option are addressed.
The report assesses the requirements, benefits, technological feasibility, and roles of Earth-to-Orbit transportation systems and options that could be developed in support of future national space programs. Transportation requirements, including those for Mission-to-Planet Earth, Space Station Freedom assembly and operation, human exploration of space, space science missions, and other major civil space missions are examined. These requirements are compared with existing, planned, and potential launch capabilities, including expendable launch vehicles (ELV's), the Space Shuttle, the National Launch System (NLS), and new launch options. In addition, the report examines propulsion systems in the context of various launch vehicles. These include the Advanced Solid Rocket Motor (ASRM), the Redesigned Solid Rocket Motor (RSRM), the Solid Rocket Motor Upgrade (SRMU), the Space Shuttle Main Engine (SSME), the Space Transportation Main Engine (STME), existing expendable launch vehicle engines, and liquid-oxygen/hydrocarbon engines. Consideration is given to systems that have been proposed to accomplish the national interests in relatively cost effective ways, with the recognition that safety and reliability contribute to cost-effectiveness. Related resources, including technology, propulsion test facilities, and manufacturing capabilities are also discussed.