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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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A popular discussion is given of microgravity effects in engineering and medicine gained from Skylab experience. Areas covered include crystal growing, liquid surface properties, diffusion, ferromagnetism, and emulsions.
A proposed NASA space research and technology development program will provide adequate data for designing closed loop life support systems for long-duration manned space missions. This program, referred to as the Pathfinder Physical-Chemical Closed Loop Life Support Program, is to identify and develop critical chemical engineering technologies for the closure of air and water loops within the spacecraft, surface habitats or mobility devices. Computerized simulation can be used both as a research and management tool. Validated models will guide the selection of the best known applicable processes and in the development of new processes. For the integration of the habitat system, a biological subsystem would be introduced to provide food production and to enhance the physical-chemical life support functions on an ever-increasing basis.
Report reviews adhesion, friction, and micromechanical properties of ceramics - properties increasingly important as more ceramic materials used in bearings, seals, and gears in advanced engines and in cutting tools and extrusion dies. Report considers effects of contaminating surface films, temperature, and chemical interactions. Examines ceramics, in both monolithic and coating form, in contact with themselves, with other harder materials, and with metals.
The Heat Flux Microsensor is a new sensor which was recently patented by Virginia Tech and is just starting to be marketed by Vatell Corp. The sensor is made using the thin-film microfabrication techniques directly on the material that is to be measured. It consists of several thin-film layers forming a differential thermopile across a thermal resistance layer. The measured heat flux q is proportional to the temperature difference across the resistance layer q= k(sub g)/delta(sub g) x (t(sub 1) - T(sub 2)), where k(sub g) is the thermal conductivity and delta (sub g) is the thickness of the thermal resistance layer. Because the gages are sputter coated directly onto the surface, their total thickness is less than 2 micrometers, which is two orders of magnitude thinner than previous gages. The resulting temperature difference across the thermal resistance layer (delta is less than 1 micrometer) is very small even at high heat fluxes. To generate a measurable signal many thermocouple pairs are put in series to form a differential thermopile. The combination of series thermocouple junctions and thin-film design creates a gage with very attractive characteristics. It is not only physically non-intrusive to the flow, but also causes minimal disruption of the surface temperature. Because it is so thin, the response time is less than 20 microsec. Consequently, the frequency response is flat from 0 to over 50 kHz. Moreover, the signal of the Heat Flux Microsensor is directly proportional to the heat flux. Therefore, it can easily be used in both steady and transient flows, and it measures both the steady and unsteady components of the surface heat flux. A version of the Heat Flux Microsensor has been developed to meet the harsh demands of combustion environments. These gages use platinum and platinum-10 percent rhodium as the thermoelectric materials. The thermal resistance layer is silicon monoxide and a protective coating of Al2O3 is deposited on top of the sensor. The superimposed thin-film pattern of all six layers is presented. The large pads are for connection with pins used to bring the signal out the back of the ceramic. In addition to the heat flux measurement, the surface temperature is measured with a platinum resistance layer (RTS). The resistance of this layer increases with increasing temperature. Therefore, these gages simultaneously measure the surface temperature and heat flux. The demonstrated applications include rocket nozzles, SCRAM jet engines, gas turbine engines, boiling heat transfer, flame experiments, basic fluid heat transfer, hypersonic flight, and shock tube testing. The laboratory involves using one of these sensors in a small combustion flame. The sensor is made on a 2.5 cm diameter piece of aluminum nitride ceramic.
The purpose of this paper is twofold: (1) to discuss high temperature superconductors with specific reference to their employment in telecommunications applications; and (2) to discuss a few of the limitations of the normally employed two-fluid model. While the debate on the actual usage of high temperature superconductors in the design of electronic and telecommunications devices - obvious advantages versus practical difficulties - needs to be settled in the near future, it is of great interest to investigate the parameters and the assumptions that will be employed in such designs. This paper deals with the issue of providing the microwave design engineer with performance data for such superconducting waveguides. The values of conductivity and surface resistance, which are the primary determining factors of a waveguide performance, are computed based on the two-fluid model. A comparison between two models - a theoretical one in terms of microscopic parameters (termed Model A) and an experimental fit in terms of macroscopic parameters (termed Model B) - shows the limitations and the resulting ambiguities of the two-fluid model at high frequencies and at temperatures close to the transition temperature. The validity of the two-fluid model is then discussed. Our preliminary results show that the electrical transport description in the normal and superconducting phases as they are formulated in the two-fluid model needs to be modified to incorporate the new and special features of high temperature superconductors. Parameters describing the waveguide performance - conductivity, surface resistance and attenuation constant - will be computed. Potential applications in communications networks and large scale integrated circuits will be discussed. Some of the ongoing work will be reported. In particular, a brief proposal is made to investigate of the effects of electromagnetic interference and the concomitant notion of electromagnetic compatibility (EMI/EMC) of high T(sub c) superconductors.
The Linear Aerospike SR-71 Experiment (LASRE) is presently being conducted to test a 20-percent-scale version of the Linear Aerospike rocket engine. This rocket engine has been chosen to power the X-33 Single Stage to Orbit Technology Demonstrator Vehicle. The rocket engine was integrated into a lifting body configuration and mounted to the upper surface of an SR-71 aircraft. This paper presents stability and control results and performance results from the envelope expansion flight tests of the LASRE configuration up to Mach 1.8 and compares the results with wind tunnel predictions. Longitudinal stability and elevator control effectiveness were well-predicted from wind tunnel tests. Zero-lift pitching moment was mispredicted transonically. Directional stability, dihedral stability, and rudder effectiveness were overpredicted. The SR-71 handling qualities were never significantly impacted as a result of the missed predictions. Performance results confirmed the large amount of wind-tunnel-predicted transonic drag for the LASRE configuration. This drag increase made the performance of the vehicle so poor that acceleration through transonic Mach numbers could not be achieved on a hot day without depleting the available fuel.
As humans extend their reach beyond the earth, bioregenerative life support systems must replace the resupply and physical/chemical systems now used. The Controlled Ecological Life Support System (CELSS) will utilize plants to recycle the carbon dioxide (CO2) and excrement produced by humans and return oxygen (O2), purified water and food. CELSS design requires knowledge of gas flux levels for net photosynthesis (PS(sub n)), dark respiration (R(sub d)) and evapotranspiration (ET). Full season gas flux data regarding these processes for wheat (Triticum aestivum), soybean (Glycine max) and rice (Oryza sativa) from published sources were used to develop empirical models. Univariate models relating crop age (days after planting) and gas flux were fit by simple regression. Models are either high order (5th to 8th) or more complex polynomials whose curves describe crop development characteristics. The models provide good estimates of gas flux maxima, but are of limited utility. To broaden the applicability, data were transformed to dimensionless or correlation formats and, again, fit by regression. Polynomials, similar to those in the initial effort, were selected as the most appropriate models. These models indicate that, within a cultivar, gas flux patterns appear remarkably similar prior to maximum flux, but exhibit considerable variation beyond this point. This suggests that more broadly applicable models of plant gas flux are feasible, but univariate models defining gas flux as a function of crop age are too simplistic. Multivariate models using CO2 and crop age were fit for PS(sub n), and R(sub d) by multiple regression. In each case, the selected model is a subset of a full third order model with all possible interactions. These models are improvements over the univariate models because they incorporate more than the single factor, crop age, as the primary variable governing gas flux. They are still limited, however, by their reliance on the other environmental conditions under which the original data were collected. Three-dimensional plots representing the response surface of each model are included. Suitability of using empirical models to generate engineering design estimates is discussed. Recommendations for the use of more complex multivariate models to increase versatility are included.
The Deep Space 4/Champollion mission is designed to test and validate technologies for landing on and anchoring to small bodies, and sample collection and transfer, in preparation for future sample return missions from comets, asteroids, and satellites. in addition, DS-4 will test technologies for advanced, multi-engine solar electric propulsion (SEP) systems, inflatable-rigidizable solar arrays, autonomous navigation and precision guidance for landing, autonomous hazard detection and avoidance, and advanced integrated avionics and packaging concepts. Deep Space-4/Champollion consists of two spacecraft: an orbiter/carrier vehicle which includes the multi-engine SEP stage, and a lander, called Champollion, which will descend to the surface of the 46P/Tempel 1 cometary nucleus. The spacecraft will launch in April, 2003 and land on the comet in September, 2006 Deep Space 4/Champollion is a joint project between NASA and CNES, the French space agency.
Analyses have been performed at the NASA Lewis Research Center's Power Systems Project Office to support the design and development of the joint U.S./Russian Solar Dynamic Flight Demonstration Project. The optical analysis of the concentrator and solar flux predictions on target receiver surfaces have an important influence on receiver design and control of the Brayton engine.
The Transition From Ignition to Flame Growth Under External Radiation in 3D (TIGER- 3D) experiment, which is slated to fly aboard the International Space Station, conducted a series of highly successful tests in collaboration with the University of Hokkaido using Japan's 10-sec JAMIC drop tower. The tests were conducted to test engineering versions of advanced flight diagnostics such as an infrared camera for detailed surface temperature measurements and an infrared spectroscopic array for gas-phase species concentrations and temperatures based on detailed spectral emissions in the near infrared. Shown in the top figure is a visible light image and in the bottom figure is an infrared image at 3.8 mm obtained during the microgravity tests. The images show flames burning across cellulose samples against a slow wind of a few centimeters per second (wind is from right to left). These flow velocities are typical of spacecraft ventilation systems that provide fresh air for the astronauts. The samples are ignited across the center with a hot wire, and the flame is allowed to spread upwind and/or downwind. As these images show, the flames prefer to spread upwind, into the fresh air, which is the exact opposite of flames on Earth, which spread much faster downwind, or with the airflow, as in forest fires.
Workshop focused on few emerging concepts(beyond tandem cells): 1. Engineering incident sun spectrum and transparency losses a) Nano emitters (dot concentrator); b) Surface plasmonics; c) Up converters; d) Down converter. 2. Intermediate band solar cells a) Efficiency projections (detail energy balance projections); b) Inserting 0,1 and 2D semiconductor structures in solar cells 3. Polymer and hybrid cells a) Nanotubes/dot polymers; b) Exciton dissociation.
In a carbothermal process, gaseous methane is passed over molten regolith, which is heated past its melting point to a temperature in excess of 1,625 C. At this temperature, materials in contact with the molten regolith (or regolith simulant) corrode and lose their structural properties. As a result, fabricating a crucible to hold the molten material and providing a method of contact heating have been problematic. Alternative containment approaches use a large crucible and limit the heat zone of the material being processed, which is inefficient because of volume and mass constraints. Alternative heating approaches use non-contact heating, such as by laser or concentrated solar energy, which can be inefficient in transferring heat and thus require higher power heat sources to accomplish processing. The innovation is a combination of materials, with a substrate material having high structural strength and stiffness and high-temperature capability, and a coating material with a high corrosion resistance and high-temperature capability. The material developed is a molybdenum substrate with an iridium coating. Creating the containment crucible or heater jacket using this material combination requires only that the molybdenum, which is easily processed by conventional methods such as milling, electric discharge machining, or forming and brazing, be fabricated into an appropriate shape, and that the iridium coating be applied to any surfaces that may come in contact with the corrosive molten material. In one engineering application, the molybdenum was fashioned into a container for a heat pipe. Since only the end of the heat pipe is used to heat the regolith, the container has a narrowing end with a nipple in which the heat pipe is snugly fit, and the external area of this nipple, which contacts the regolith to transfer heat into it, is coated with iridium. At the time of this reporting, no single material has been found that can perform the functions of this combination of materials, and other combinations of materials have not proven to be survivable to the corrosiveness of this environment. High-temperature processing of materials with similar constituencies as lunar regolith is fairly common. The carbo-thermal process is commonly used to make metallurgical-grade silicon for the semiconductor and solar-cell industries.
This report will address materials selection for components in a proposed Venus lander system. The lander would use active refrigeration to allow Space Science instrumentation to survive the extreme environment that exists on the surface of Venus. The refrigeration system would be powered by a Stirling engine-based system and is termed the Advanced Stirling Duplex (ASD) concept. Stirling engine power conversion in its simplest definition converts heat from radioactive decay into electricity. Detailed design decisions will require iterations between component geometries, materials selection, system output, and tolerable risk. This study reviews potential component requirements against known materials performance. A lower risk, evolutionary advance in heater head materials could be offered by nickel-base superalloy single crystals, with expected capability of approximately 1100C. However, the high temperature requirements of the Venus mission may force the selection of ceramics or refractory metals, which are more developmental in nature and may not have a well-developed database or a mature supporting technology base such as fabrication and joining methods.
Identification of innovative solutions to unique materials problems is an every-day quest for members of the aerospace community. Finding a technique that will minimize costs, maximize throughput, and generate quality results is always the target. United Space Alliance Materials Engineers recently conducted such a search in their drive to return the Space Shuttle fleet to operational status. The removal of high performance thermal coatings from solid rocket motors represents a formidable task during post flight disassembly on reusable expended hardware. The removal of these coatings from unfired motors increases the complexity and safety requirements while reducing the available facilities and approved processes. A temporary solution to this problem was identified, tested and approved during the Solid Rocket Booster (SRB) return to flight activities. Utilization of ultra high-pressure liquid nitrogen (LN2) to strip the protective coating from assembled space shuttle hardware marked the first such use of the technology in the aerospace industry. This process provides a configurable stream of liquid nitrogen (LN2) at pressures of up to 55,000 psig. The performance of a one-time certification for the removal of thermal ablatives from SRB hardware involved extensive testing to ensure adequate material removal without causing undesirable damage to the residual materials or aluminum substrates. Testing to establish appropriate process parameters such as flow, temperature and pressures of the liquid nitrogen stream provided an initial benchmark for process testing. Equipped with these initial parameters engineers were then able to establish more detailed test criteria that set the process limits. Quantifying the potential for aluminum hardware damage represented the greatest hurdle for satisfying engineers as to the safety of this process. Extensive testing for aluminum erosion, surface profiling, and substrate weight loss was performed. This successful project clearly demonstrated that the liquid nitrogen jet possesses unique strengths that align remarkably well with the unusual challenges that space hardware and missile manufacturers face on a regular basis. Performance of this task within the confines of a critical manufacturing facility marks a milestone in advanced processing.
The Cassini spacecraft was launched on October 15, 1997 and arrived at Saturn on June 30, 2004. It has performed detailed observations and remote sensing of Saturn, its rings, and its satellites since that time. Cassini deployed the European-built Huygens probe which descended through the Titan atmosphere and landed on its surface on January 14, 2005. Operating the Cassini spacecraft is a complex scientific, engineering, and management job. In order to safely operate the spacecraft, a large number of flight rules were developed. These flight rules must be enforced throughout the lifetime of the Cassini spacecraft. Flight rules are defined as any operational limitation imposed by the spacecraft system design, hardware, and software, violation of which would result in spacecraft damage, loss of consumables, loss of mission objectives, loss and/or degradation of science, and less than optimal performance. Flight rules require clear description and rationale. Detailed automated methods have been developed to insure the spacecraft is continuously operated within these flight rules. An overview of all the flight rules allocated to the Cassini Attitude Control and Articulation Subsystem and how they are enforced is presented in this paper.
With increased emphasis on lunar exploration and scientific investigation, there is a desire to deliver a wide variety of payloads to the lunar surface. Many of these payloads will require the use of surface mobility capability such as a rover. NASA has combined spacecraft and subsystem engineers from across the Agency to develop a ‘pallet’ lander design intended to deliver and easily deploy a medium-sized payload (~300 kg) to the polar regions of the Moon. The lander provides power to the payload from transit soon after lunar landing. The lander is not intended to survive the lunar night. The design of the lander was based on a minimum set of level 1 requirements where traditional risk, mass, and performance trade parameters were weighed lower than cost. In other words, the team did not sacrifice ‘good enough’ for ‘better’ or ‘best.’ As a NASA class D spacecraft (as defined in NPR 8705.4, Risk Classification for NASA Payloads, the lander employs single-string (i.e., zero-fault-tolerant) systems as a baseline. The design utilizes existing technologies and components where possible, though some enhancements have been targeted in areas such as precision autonomous landing and low-cost structural design/fabrication. It is important to note that these and other derived technologies are extensible to other lander designs and missions. This TP describes the requirements and approaches upon which the lander design is based; discusses key design decisions, analyses, and trades used to derive the design; provides a snapshot of each major subsystem; and identifies open items, issues, and challenges for which work is continuing.
A trade space exploration of a new NASA regional transport aircraft concept called the SUbsonic Single Aft eNgine (SUSAN) Electrofan is presented. The SUSAN concept uses a 20MW Electrified Aircraft Propulsion (EAP) system to enable advance Propulsion Airframe Integration (PAI) in transport category aircraft. Alternative fuels will be used to reduce the amount of emissions per energy used. By combining these features there is the potential to reduce aircraft emissions by 50% per passenger/mile while retaining the size, speed, and range of large regional jets. SUSAN is has a 750 mile economic mission, a 2500 mile design range and a maximum capacity of 180 passengers. The SUSAN configuration utilizes a single aft mounted engine and distributed electric wing-mounted thrusters on a tube and wing arrangement with a T-tail empennage. The SUSAN Electrofan employs a hybrid powertrain to enable: single turbofan operation on a large transport category aircraft; increased aerodynamic and propulsive efficiency through placement of electric engines; optimized turbofan sizing and efficiency through control and electric boosting, reduced control surface sizing through thrust augmentation. A single use battery is employed as the power source in case of turbofan failure. The design study also considers the constraints of operating within the current airport, airspace and economic constraints. This paper presents the status of the trade space exploration; however the concept definition is not finished. Forward work includes optimizing the overall aircraft configuration and including certain hard to model features like boundary layer ingestion or natural laminar flow across all appliable subsystems. Additional work forward work is a more extensive analysis of the configuration using alternative fuels.
The SUSAN concept uses a 20-megawatt Electrified Aircraft Propulsion system to enable advance Propulsion Airframe Integration in transport category aircraft. Alternative fuels will be used to reduce the amount of emissions per energy used. By combining these features there is the potential to reduce aircraft emissions by 50 percent per passenger/mile while retaining the size, speed, and range of large regional jets. The SUSAN is has a 750-mile economic mission, a 2500-mile design range and a maximum capacity of 180 passengers. The SUSAN configuration utilizes a single aft mounted engine and distributed electric wing-mounted thrusters on a tube and wing arrangement with a T-tail empennage. The SUSAN Electrofan employs a hybrid powertrain to enable: single turbofan operation on a large transport category aircraft; increased aerodynamic and propulsive efficiency through placement of electric engines; optimized turbofan sizing and efficiency through control and electric boosting, reduced control surface sizing through thrust augmentation. A single use battery is employed as the power source in case of turbofan failure. The design study also considers the constraints of operating within the current airport, airspace, and economic constraints. Forward work includes optimizing the overall aircraft configuration and including certain hard to model features like boundary layer ingestion or natural laminar flow across all appliable subsystems. Additional work forward work is a more extensive analysis of the configuration using alternative fuels.
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