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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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At least 235 records · Page 13
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NASA NTRS2004

Mid-frequency Band Dynamics of Large Space Structures

High and low intensity dynamic environments experienced by a spacecraft during launch and on-orbit operations, respectively, induce structural loads and motions, which are difficult to reliably predict. Structural dynamics in low- and mid-frequency bands are sensitive to component interface uncertainty and non-linearity as evidenced in laboratory testing and flight operations. Analytical tools for prediction of linear system response are not necessarily adequate for reliable prediction of mid-frequency band dynamics and analysis of measured laboratory and flight data. A new MATLAB toolbox, designed to address the key challenges of mid-frequency band dynamics, is introduced in this paper. Finite-element models of major subassemblies are defined following rational frequency-wavelength guidelines. For computational efficiency, these subassemblies are described as linear, component mode models. The complete structural system model is composed of component mode subassemblies and linear or non-linear joint descriptions. Computation and display of structural dynamic responses are accomplished employing well-established, stable numerical methods, modern signal processing procedures and descriptive graphical tools. Parametric sensitivity and Monte-Carlo based system identification tools are used to reconcile models with experimental data and investigate the effects of uncertainties. Models and dynamic responses are exported for employment in applications, such as detailed structural integrity and mechanical-optical-control performance analyses.

dynamics↗
NASA NTRSDate not supplied

Structural Requirements and Scaling Analysis of a Fluidic Mirror Space Telescope Support Structure

The NASA FLUTE project proposes large-scale (50m) fluidic telescopes for astronomy applications. To continue to explore the universe, astronomers require larger and larger telescope apertures. The highest priority astrophysics targets such as exoplanets and early galaxies are extremely faint, motivating larger telescope apertures. However, mission costs depend on aperture diameter, and scaling apertures beyond 10-m apertures faces economic and technological viability challenges. An unsegmented primary mirror made in space via fluidic microgravity shaping would provide a scalable and cost-effective method to scale apertures to 50-m scale while achieving sub-nanometer (root mean square) surface quality. Such microgravity fluidic shaping has been demonstrated in laboratory neutral buoyance environments, parabolic microgravity experiments, as well aboard the International Space Station. One of the main components of a fluidic observatory is the mirror frame. The frame must provide a stable bounding circular ring which the edges of the fluid mirror surface can wet. The frame can optionally provide a ‘floor’ surface on the interior of the ring to provide additional fluid support and reduce required fluid volume. In this work, we evaluate several classes of structural frame architectures potentially suitable for a fluidic telescope support structure. We start by estimating stability requirements, orbital, station keeping, and slew loads based on a notional CONOPS. The scaling of overall fluid mass required for each architecture is evaluated. Preliminary results elucidate the importance of a support floor for overall mission viability above 10-m diameter. We then investigate the scaling of a tetrahedral truss frame support structure. We show that segmented solid shell support surfaces can provide sufficient stability at modest mass fractions. We estimate that the total fluid and frame mass for a 50-m telescope could be on the order of 15,000 kg. Finally, implementation considerations are discussed, including deployment/assembly methodologies. The results of this study establish feasibility of a large-scale fluidic telescope and will guide further architecture development and detailed structural design.

Christine Gregg↗
NASA NTRS1992

Method of delivering lunar generated fluid to earth orbit using an external tank

A method and apparatus are provided for delivering lunar generated fluid to Earth orbit from lunar orbit. Transport takes place in an external tank of a shuttle which has been suitably outfitted in Earth orbit for reusable travel between Earth orbit and a lunar orbit. The outfitting of the external tank includes the adding of an engine, an electrical system, a communication system, a guidance system, an aerobraking device, and a plurality of interconnected fluid storage tanks to the hydrogen and oxygen tanks of the external tank. The external tank is then propelled to lunar orbit the first time using Earth-based propellant. In lunar orbit, the storage tanks are filled with the lunar generated fluid with the remainder tank volumes filled with lunar generated liquid oxygen and hydrogen which serve as propellants for returning the tank to Earth orbit where the fluid is off-loaded. The remaining lunar generated oxygen and hydrogen is then sufficient to return the external tank to lunar orbit so that a subsequent cycle of fluid delivery is repeated. A space station in a higher Earth orbit is preferably used to outfit the external tank, and a lunar node in lunar orbit is used to store and transfer the fluid and liquid oxygen and hydrogen to the external tank. The lunar generated fluid is preferably .sup.3 He.

Butterfield, Ansel J.↗
NASA NTRS2006

Shuttle Debris Impact Tool Assessment Using the Modern Design of Experiments

Computational tools have been developed to estimate thermal and mechanical reentry loads experienced by the Space Shuttle Orbiter as the result of cavities in the Thermal Protection System (TPS). Such cavities can be caused by impact from ice or insulating foam debris shed from the External Tank (ET) on liftoff. The reentry loads depend on cavity geometry and certain Shuttle state variables, among other factors. Certain simplifying assumptions have been made in the tool development about the cavity geometry variables. For example, the cavities are all modeled as shoeboxes , with rectangular cross-sections and planar walls. So an actual cavity is typically approximated with an idealized cavity described in terms of its length, width, and depth, as well as its entry angle, exit angle, and side angles (assumed to be the same for both sides). As part of a comprehensive assessment of the uncertainty in reentry loads estimated by the debris impact assessment tools, an effort has been initiated to quantify the component of the uncertainty that is due to imperfect geometry specifications for the debris impact cavities. The approach is to compute predicted loads for a set of geometry factor combinations sufficient to develop polynomial approximations to the complex, nonparametric underlying computational models. Such polynomial models are continuous and feature estimable, continuous derivatives, conditions that facilitate the propagation of independent variable errors. As an additional benefit, once the polynomial models have been developed, they require fewer computational resources to execute than the underlying finite element and computational fluid dynamics codes, and can generate reentry loads estimates in significantly less time. This provides a practical screening capability, in which a large number of debris impact cavities can be quickly classified either as harmless, or subject to additional analysis with the more comprehensive underlying computational tools. The polynomial models also provide useful insights into the sensitivity of reentry loads to various cavity geometry variables, and reveal complex interactions among those variables that indicate how the sensitivity of one variable depends on the level of one or more other variables. For example, the effect of cavity length on certain reentry loads depends on the depth of the cavity. Such interactions are clearly displayed in the polynomial response models.

DeLoach, R.↗
NASA NTRS1991

Parametric entry corridors for lunar/Mars aerocapture missions

Parametric atmospheric entry corridor data are presented for Earth and Mars aerocapture. Parameter ranges were dictated by the range of mission designs currently envisioned as possibilities for the Human Exploration Initiative (HEI). This data, while not providing a means for exhaustive evaluation of aerocapture performance, should prove to be a useful aid for preliminary mission design and evaluation. Entry corridors are expressed as ranges of allowable vacuum periapse altitude of the planetary approach hyperbolic orbit, with chart provided for conversion to an approximate flight path angle corridor at entry interface (125 km altitude). The corridor boundaries are defined by open-loop aerocapture trajectories which satisfy boundary constraints while utilizing the full aerodynamic control capability of the vehicle (i.e., full lift-up or full lift-down). Parameters examined were limited to those of greatest importance from an aerocapture performance standpoint, including the approach orbit hyperbolic excess velocity, the vehicle lift to drag ratio, maximum aerodynamic load factor limit, and the apoapse of the target orbit. The impact of the atmospheric density bias uncertainties are also included. The corridor data is presented in graphical format, and examples of the utilization of these graphs for mission design and evaluation are included.

Ling, Lisa M.↗
NASA NTRS1994

GSOSTATS Database: USAF Synchronous Satellite Catalog Data Conversion Software. User's Guide and Software Maintenance Manual, Version 2.1

The United States Air Force (USAF) provides NASA Lewis Research Center with monthly reports containing the Synchronous Satellite Catalog and the associated Two Line Mean Element Sets. The USAF Synchronous Satellite Catalog supplies satellite orbital parameters collected by an automated monitoring system and provided to Lewis Research Center as text files on magnetic tape. Software was developed to facilitate automated formatting, data normalization, cross-referencing, and error correction of Synchronous Satellite Catalog files before loading into the NASA Geosynchronous Satellite Orbital Statistics Database System (GSOSTATS). This document contains the User's Guide and Software Maintenance Manual with information necessary for installation, initialization, start-up, operation, error recovery, and termination of the software application. It also contains implementation details, modification aids, and software source code adaptations for use in future revisions.

Mallasch, Paul G.↗
NASA NTRS2005

Flight Loads and Environments Initiative

A viewgraph presentation on the design of a lightweight non-intrusive force measurement device (FMD) to reduce the cost per effective payload (PL) mass into orbit (CPMO) by improving launch vehicle (LV) loads and environments.

Kaufman, Daniel↗
NASA NTRS2009

Development, Testing, and Failure Mechanisms of a Replicative Ice Phase Change Material Heat Exchanger

Phase change materials (PCM) may be useful for thermal control systems that involve cyclical heat loads or cyclical thermal environments such as Low Earth Orbit (LEO) and Low Lunar Orbit (LLO). Thermal energy can be stored in the PCM during peak heat loads or in adverse thermal environments. The stored thermal energy can then be released later during minimum heat loads or in more favorable thermal environments. One advantage that PCM s have over evaporators in this scenario is that they do not use a consumable. Wax PCM units have been baselined for the Orion thermal control system and also provide risk mitigation for the Altair Lander. However, the use of water as a PCM has the potential for significant mass reduction since the latent heat of formation of water is approximately 70% greater than that of wax. One of the potential drawbacks of using ice as a PCM is its potential to rupture its container as water expands upon freezing. In order to develop a space qualified ice PCM heat exchanger, failure mechanisms must first be understood. Therefore, a methodical experimental investigation has been undertaken to demonstrate and document specific failure mechanisms due to ice expansion in the PCM. An ice PCM heat exchanger that replicates the thermal energy storage capacity of an existing wax PCM unit was fabricated and tested. Additionally, methods for controlling void location in order to reduce the risk of damage due to ice expansion are investigated. This paper presents the results to date of this investigation.

Leimkuehler, Thomas O.↗
NASA NTRS2010

Development, Testing, and Failure Mechanisms of a Replicative Ice Phase Change Material Heat Exchanger

Phase change materials (PCM) may be useful for thermal control systems that involve cyclical heat loads or cyclical thermal environments such as Low Earth Orbit (LEO) and Low Lunar Orbit (LLO). Thermal energy can be stored in the PCM during peak heat loads or in adverse thermal environments. The stored thermal energy can then be released later during minimum heat loads or in more favorable thermal environments. One advantage that PCM's have over evaporators in this scenario is that they do not use a consumable. Wax PCM units have been baselined for the Orion thermal control system and also provide risk mitigation for the Altair Lander. However, the use of water as a PCM has the potential for significant mass reduction since the latent heat of formation of water is approximately 70% greater than that of wax. One of the potential drawbacks of using ice as a PCM is its potential to rupture its container as water expands upon freezing. In order to develop a space qualified ice PCM heat exchanger, failure mechanisms must first be understood. Therefore, a methodical experimental investigation has been undertaken to demonstrate and document specific failure mechanisms due to ice expansion in the PCM. An ice PCM heat exchanger that replicates the thermal energy storage capacity of an existing wax PCM unit was fabricated and tested. Additionally, methods for controlling void location in order to reduce the risk of damage due to ice expansion are investigated. This paper presents the results to date of this investigation. Nomenclature

Leimkuehler, Thomas O.↗
NASA NTRS2010

Testing and Failure Mechanisms of Ice Phase Change Material Heat Exchangers

Phase change materials (PCM) may be useful for thermal control systems that involve cyclical heat loads or cyclical thermal environments such as Low Earth Orbit (LEO) and Low Lunar Orbit (LLO). Thermal energy can be stored in the PCM during peak heat loads or in adverse thermal environments. The stored thermal energy can then be released later during minimum heat loads or in more favorable thermal environments. One advantage that PCM s have over evaporators in this scenario is that they do not use a consumable. The use of water as a PCM rather than the more traditional paraffin wax has the potential for significant mass reduction since the latent heat of formation of water is approximately 70% greater than that of wax. One of the potential drawbacks of using ice as a PCM is its potential to rupture its container as water expands upon freezing. In order to develop a space qualified ice PCM heat exchanger, failure mechanisms must first be understood. Therefore, a methodical experimental investigation has been undertaken to demonstrate and document specific failure mechanisms due to ice expansion in the PCM. A number of ice PCM heat exchangers were fabricated and tested. Additionally, methods for controlling void location in order to reduce the risk of damage due to ice expansion were investigated. This paper presents the results of testing that occurred from March through September of 2010 and builds on testing that occurred during the previous year.

Leimkuehler, Thomas O.↗
NASA NTRS1990

The effects of the space environment on two aramid materials

Two aramid fibers having closely related chemical structures were chosen for important roles in the first tether to be used to connect pairs of orbiting vehicles. The protective outer jackets of the tethers will consist of woven fibers of poly(m-phenylene isophthalamide), commercially available from du Pont as Nomex. A cylindrical sheath of woven Kevlar 29, whose principal constituent is poly(p-phenylene terephthalamide), will be the load-bearing component for the tethers. Orbiting tethers will be in a hostile environment in which short wavelength electromagnetic radiation and energetic charged particles degrade exposed organic materials. At lower orbiting altitudes atomic oxygen is an especially serious hazard. Studies on the effects of ultraviolet radiation and atomic oxygen on fibers and films of Kevlar and Nomex are in progress. In an experiment to simulate the effects of atomic oxygen in space, small tows of Kevlar and Nomex were mounted in a commercial ashing device filled with oxygen at low pressure. An RF discharge in the instrument dissociated the molecular oxygen producing a strongly oxidizing atmosphere containing O(3P)(sup 4). Erosion was measured in terms of mass loss. Kevlar films were exposed to UV radiation in an apparatus consisting of a small vacuum chamber, 23 cm in diameter, into which a mass spectrometer and a quartz window were incorporated. Samples were exposed under vacuum with a 1000 watt xenon-arc lamp. Volatile products could be monitored with the mass spectrometer during the exposures. Transmission infrared spectra were taken before and after exposure to monitor chemical changes in the films.

Kiefer, Richard L.↗
NASA NTRS1976

Results of tests CS4 and CS5 to investigate dynamic loads and pressures on 0.03-scale models (Ax1319-3/4 and 45-0) of mated 747 cam and space shuttle orbiter in the Boeing transonic wind tunnel

A 0.03-scale model of the 747 CAM/Orbiter was tested in an 8 x 12 foot transonic wind tunnel. Dynamic loads, pressure, and empennage flow field data were obtained using pressure transducers, strain gages, and a split film anemometer. The test variables included Mach number, angle of attack, sideslip angle, orbiter tailcone on and off, orbiter partial tailcone, orbiter nozzle air scoops, orbiter body flap angle, and orbiter elevon angle.

Source record↗
NASA NTRS1985

Space shuttle orbiter auxiliary power unit development challenges

When the flying spacecraft was approved for development, a power unit for the hydraulic system had to be developed. Unlike other systems on the orbiter, there was no precedent in earlier spacecraft for a hydraulic system nor for the power unit to drive the hydraulic pumps. The only prototypes available were airplane auxiliary power units (APU), which were not required to operate in the severe environments of a spacecraft nor to have the longevity of an orbiter hydraulic power unit. The challenge was to build a hydraulic power unit which could operate in 0g or 3g, in a vacuum or at sea level pressure, and at -65 F or 225 F, which would be capable of restarting while hot, and which would be capable of sustaining the hydraulic loads for the life of the orbiter. The basic approach to providing hydraulic power for the orbiter was to use a small, high speed, monopropellant fueled turbine power unit to drive a conventional aircraft type hydraulic pump. The stringent requirements imposed on the orbiter APU quickly made this machine different from existing aircraft APUs.

Lance, R.↗
NASA NTRS1984

Integration and flight demonstration of a high-capacity monogroove heat-pipe radiator

The cancellation of the TDRS-B satellite as the payload for the eighth Space Shuttle mission provided a unique opportunity to demonstrate on-orbit operation of the high-capacity monogroove heat pipe used in the space constructible radiator subsystem. In less than 4 months, a flight experiment was conceived, designed, fabricated, tested, integrated with a payload carrier, installed in the Orbiter Challenger payload bay, and successfully operated in flight. Still color photographs and direct crew visual observation of color changes in a pattern of temperature-sensitive liquid-crystal tapes provided the temperature data necessary to verify successful on-orbit startup and orbital transient response of the heat pipe when subjected to a heat load from its attached electrical heaters. This successful on-orbit demonstration verified analytical design tools and provided confidence in the use of high-capacity heat pipes for future space applications. The flight experiment hardware and the integration and test activities that led to the flight are described, and the actual flight results are compared to analytical performance predictions.

Rankin, J. G.↗