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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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Viewgraphs and discussion on advanced modular power supplies for Space Station Freedom are presented. Topics covered include concept and characteristics; user power supply applications; and bulk converter application.
An instrumentation, called Space Active Modular Materials ExperimentS (SAMMES), is developed for monitoring the spacecraft environment and for accurately measuring the degradation of space materials in low earth orbit (LEO). The SAMMES architecture concept can be extended to instrumentation for planetary exploration, both on spacecraft and in situ. The operating environment for planetary application will be substantially different, with temperature extremes and harsh solar wind and cosmic ray flux on lunar surfaces and temperature extremes and high winds on venusian and Martian surfaces. Moreover, instruments for surface deployment, which will be packaged in a small lander/rover (as in MESUR, for example), must be extremely compact with ultralow power and weight. With these requirements in mind, the SAMMES concept was extended to a sensor/instrumentation scheme for the lunar and Martian surface environment.
Spaceflight optical instruments have two conflicting requirements. They need to be both rigid and lightweight. In addition, for interferometric far ultraviolet spectrometers, the requirements for precision positioning are more severe than for conventional spectrometers. To meet the challenge of lightweight optical instruments, a modular adjustment mechanism was developed to position two orthogonal axes of a universal three-axis gimbal support system with a positioning accuracy on the order of 10 arc seconds. The mechanism was designed as a self-contained assembly which can be removed after final alignment of a spacebound optical instrument to reduce its in-flight mass. To demonstrate the concept, a number of these assemblies were made and mounted on two of the positioning axes of a far ultraviolet spatial heterodyne interferometer. A shaft clamp was used on each positioning axis to retain the adjusted position. This paper describes the design of the mechanism and presents optical test results.
An extension is proposed for the NASA Space Exploration Initiative's Generic Modular Flow Schematics for physical/chemical life support systems which involves the addition of biological processes. The new system architecture includes plant, microbial, and animal habitat, as well as the human habitat subsystem. Major Feedstock Production and Food Preparation and Packaging components have also been incorporated. Inedible plant, aquaculture, microbial, and animal solids are processed for recycling.
A study was performed to evaluate the reliability of Integrated Modular Engine (IME) concepts. Comparisons were made between networked IME systems and non-networked discrete systems using expander cycle configurations. Both redundant and non-redundant systems were analyzed. Binomial approximation and Markov analysis techniques were employed to evaluate total system reliability. In addition, Failure Modes and Effects Analyses (FMEA), Preliminary Hazard Analyses (PHA), and Fault Tree Analysis (FTA) were performed to allow detailed evaluation of the IME concept. A discussion of these system reliability concepts is also presented.
A study was performed to develop a fluid system design and show the feasibility of constructing an integrated modular engine (IME) configuration, using an expander cycle engine. The primary design goal of the IME configuration was to improve the propulsion system reliability. The IME fluid system was designed as a single fault tolerant system, while minimizing the required fluid components. This study addresses the design of the high pressure manifolds, turbopumps and thrust chambers for the IME configuration. A physical layout drawing was made, which located each of the fluid system components, manifolds and thrust chambers. Finally, a comparison was made between the fluid system designs of an IME system and a non-network (clustered) engine system.
Space vehicle propulsion systems are traditionally comprised of a cluster of discrete engines, each with its own set of turbopumps, valves, and a thrust chamber. The Integrated Modular Engine (IME) concept proposes a vehicle propulsion system comprised of multiple turbopumps, valves, and thrust chambers which are all interconnected. The IME concept has potential advantages in fault-tolerance, weight, and operational efficiency compared with the traditional clustered engine configuration. The purpose of this study is to examine the steady-state performance of an IME system with various components removed to simulate fault conditions. An IME configuration for a hydrogen/oxygen expander cycle propulsion system with four sets of turbopumps and eight thrust chambers has been modeled using the Rocket Engine Transient Simulator program. The nominal steady-state performance is simulated, as well as turbopump, thrust chamber, and duct failures. The impact of component failures on system performance is discussed in the context of the system's fault tolerant capabilities.
The paper proposes a modular approach to rocket propulsion which offers a versatile method for realizing the goals of low cost, safety, reliability, and ease of operation. It is shown that, using practical modules made up of only 4-6 individual elements, it is possible to achieve thrust levels of 2-3 mln lbf and more, using turbomachinery, thrust chambers, lines, and valves about the size of SSME hardware. The approach is illustrated by a LOX/LH2 configuration.
Three modular heat-transfer components designed for use together or separately. Simple mechanical connections facilitate assembly of these and related heat-transfer components into cooling systems of various configurations, such as to cool laboratory equipment rearranged for different experiments. Components are clamp-on cold plate, cold plate attached to flexible heat pipe, and thermal-bus receptacle. Clamp-on cold plate moved to any convenient location for attachment of equipment cooled by it, then clamped onto thermal bus. Heat from equipment conducted through plate and into coolant. Thermal-bus receptacle integral with thermal bus. Includes part of thermal bus to which clamp-on cold plate attached, plus tapered socket into which condenser end of flexible heat pipe plugged. Thermal-bus receptacle includes heat-pipe wick structure using coolant in bus to enhance transfer of heat from cold plate.
N-Modular Redundancy (NMR) is one of the best known fault tolerance techniques. Replication of a module to achieve fault tolerance is in some ways analogous to the use of a repetition code where an information symbol is replicated as parity symbols in a codeword. Linear Error-Correcting Codes (ECC) use linear combinations of information symbols as parity symbols which are used to generate syndromes for error patterns. These observations indicate links between the theory of ECC and the use of hardware redundancy for fault tolerance. In this paper, we explore some of these links and show examples of NMR systems where identification of good and failed elements is accomplished in a manner similar to error correction using linear ECC's.
This paper describes a Modular Artificial Intelligence Inference Engine System (MAIS) support tool that would provide health and status monitoring, cognitive replanning, analysis and support of on-orbit Space Station, Spacelab experiments and systems.
Space vehicle propulsion systems are traditionally comprised of a cluster of discrete engines, each with its own set of turbopumps, valves, and a thrust chamber. The Integrated Modular Engine (IME) concept proposes a vehicle propulsion system comprised of multiple turbopumps, valves, and thrust chambers which are all interconnected. The IME concept has potential advantages in fault-tolerance, weight, and operational efficiency compared with the traditional clustered engine configuration. The purpose of this study is to examine the steady-state performance of an IME system with various components removed to simulate fault conditions. An IME configuration for a hydrogen/oxygen expander cycle propulsion system with four sets of turbopumps and eight thrust chambers has been modeled using the Rocket Engine Transient Simulator (ROCETS) program. The nominal steady-state performance is simulated, as well as turbopump thrust chamber and duct failures. The impact of component failures on system performance is discussed in the context of the system's fault tolerant capabilities.
Conceptual system for positioning mobile robotic manipulators on large main structure includes modular tracks and ancillary structures assembled easily along with main structure. System, called "tracked robotic location system" (TROLS), originally intended for application to platforms in outer space, but TROLS concept might also prove useful on Earth; for example, to position robots in factories and warehouses. T-cross-section rail keeps mobile robot on track. Bar codes mark locations along track. Each robot equipped with bar-code-recognizing circuitry so it quickly finds way to assigned location.
The Large Area Modular Array of Reflectors (LAMAR) was selected as an attached payload for the Space Station Freedom in June 1989. It is a high throughput multiple X-ray telescope system that was capable of performing both imaging and dispersive spectroscopy measurements simultaneously. The LAMAR program was halted in 1990 when attached payloads were removed from the Space Station. A technical description of LAMAR is the main body of this report.
The effective performance of modular thrusters in an aerospike configuration is difficult to determine. Standard analytical tools are applicable to conventional nozzle shapes, but are limited when applied to an aerospike nozzle (An aerospike nozzle is an altitude compensating external nozzle). Three baseline nozzle shapes are derived using standard analytical procedures. The baseline nozzle sizes are restricted to fill a volume envelope. The three shapes are an axi-symmetric round nozzle, a two dimensional planar square exit nozzle, and a super elliptic round to nearly square nozzle. The integrated (thruster/aerospike) performance of the three nozzles is determined through the use of three dimensional viscous computational fluid dynamic (CFD) calculation where complex features of the flow field can be accurately captured. The resulting installed performance is then used to evaluate the efficiency of these nozzle shapes for aerospike applications. The determination of effective performance of a thruster nozzle integrated into an aerospike nozzle require the solution of the three dimensional turbulent Navier-Stokes equations. The model used in this study consisted of two zones; one of the upstream thruster cowl surface so freestream conditions can be accurately predicted, and two, the aerospike surface beginning with the thruster outflow and extending to the end of the aerospike surface. The numerical grid consisted of over 120,000 nodes and used symmetry on the thruster centerline and edge. A two species non-reacting chemistry model was used to capture the variation of fluid properties between the hot plume base and freestream air. From the results of the three baseline nozzle aerospike calculations, the effictive performance of the nozzle was determined. The flow fields of these calculations do show some variation between the cases. Recirculation zones on the cowl surface is predicted for the two dimensional planar nozzle and a smaller one for the super elliptic nozzle. The recirculation is caused by the strong pressure gradient between the plume and freestream flows. The axi-symmetric nozzle results indicate recirculation zones on the thruster face. These recirculation zones smooth the pressure gradient between the plume and freestream flow limiting the formation of recirculation on the cowl surface. Thruster to thruster interaction is evident for the axi-symmetric and supper elliptic calculation while the two dimensional planar nozzle did not have any lateral expansion in the nozzle, so thruster to thruster interaction is limited. The integrated performance results, at the altitude choosen, show very little variation between the three thruster shapes. This result allows for nozzle shape determination based on additional considerations (thermal, structural, weight) besides performance.
The Modular Manufacturing Simulator (MMS) has been developed for the beginning user of computer simulations. Consequently, the MMS cannot model complex systems that require branching and convergence logic. Once a user becomes more proficient in computer simulation and wants to add more complexity, the user is encouraged to use one of the many available commercial simulation systems. The (MMS) is based on the SSE5 that was developed in the early 1990's by the University of Alabama in Huntsville (UAH). A recent survey by MSFC indicated that the simulator has been a major contributor to the economic impact of the MSFC technology transfer program. Many manufacturers have requested additional features for the SSE5. Consequently, the following features have been added to the MMS that are not available in the SSE5: runs under Windows, print option for both input parameters and output statistics, operator can be fixed at a station or assigned to a group of stations, operator movement based on time limit, part limit, or work-in-process (WIP) limit at next station. The movement options for a moveable operators are: go to station with largest WIP, rabbit chase where operator moves in circular sequence between stations, and push/pull where operator moves back and forth between stations. This user's manual contains the necessary information for installing the MMS on a PC, a description of the various MMS commands, and the solutions to a number of sample problems using the MMS. Also included in the beginning of this report is a brief discussion of technology transfer.
The aviation industry is gradually moving toward the use of integrated modular avionics (IMA) for civilian transport aircraft. An important concern for IMA is ensuring that applications are safely partitioned so they cannot interfere with one another. We have investigated the problem of ensuring safe partitioning and logical non-interference among separate applications running on a shared Avionics Computer Resource (ACR). This research was performed in the context of ongoing standardization efforts, in particular, the work of RTCA committee SC-182, and the recently completed ARINC 653 application executive (APEX) interface standard. We have developed a formal model of partitioning suitable for evaluating the design of an ACR. The model draws from the mathematical modeling techniques developed by the computer security community. This report presents a formulation of partitioning requirements expressed first using conventional mathematical notation, then formalized using the language of SRI'S Prototype Verification System (PVS). The approach is demonstrated on three candidate designs, each an abstraction of features found in real systems.
A comparison of space experiments with previous missions shows a common theme. Some of the recent experiments are based on the scientific fundamentals of instruments of prior years. However, the main distinguishing characteristic is the embodiment of advances in engineering and manufacturing in order to extract clearer and sharper images and extend the limits of measurement. One area of importance to future missions is providing vibration free observation platforms at acceptable costs. It has been shown by researchers that vibration problems cannot be eliminated by passive isolation techniques alone. Therefore, various organizations have conducted research in the area of combining active and passive vibration control techniques. The essence of this paper is to present progress in what is believed to be a new concept in this arena. It is based on the notion that if one active element in a vibration transmission path can provide a reasonable vibration attenuation, two active elements in series may provide more control options and better results. The paper presents the functions of a modular split shaft linear actuator developed by NASA's Goddard Space Flight Center and University of Massachusetts Lowell. It discusses some of the control possibilities facilitated by the device. Some preliminary findings and problems are also discussed.