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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 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS2005

Dynamic Weakening (Extinction) of Simple Hydrocarbon-air Counterflow Diffusion Flames by Oscillatory Inflows

This study of laminar non-premixed HC-air flames used an Oscillatory-input Opposed Jet Burner (OOJB) system developed from a previously well-characterized 7.2-mm Pyrex-nozzle OJB system. Over 600 dynamic Flame Strength (FS) measurements were obtained on unanchored (free-floating) laminar Counterflow Diffusion Flames (CFDF's). Flames were stabilized using plug inflows having steady-plus-sinusoidal axial velocities of varied magnitude, frequency, f, up to 1600 Hz, and phase angle from 0 (most data) to 360 degrees. Dynamic FS is defined as the maximum average air input velocity (U(sub air), at nozzle exit) a CFDF can sustain before strain-induced extinction occurs due to prescribed oscillatory peak-to-peak velocity inputs superimposed on steady inputs. Initially, dynamic flame extinction data were obtained at low f, and were supported by 25-120 Hz Hot-Wire cold-flow velocity data at nozzle exits. Later, expanded extinction data were supported by 4-1600 Hz Probe Microphone (PM) pk/pk P data at nozzle exits. The PM data were first obtained without flows, and later with cold stagnating flows, which better represent speaker-diaphragm dynamics during runs. The PM approach enabled characterizations of Dynamic Flame Weakening (DFW) of CFDF's from 8 to 1600 Hz. DFW was defined as % decrease in FS per Pascal of pk/pk P oscillation, namely, DFW = - 100 d(U(sub air) / U(sub air),0Hz) / d(pkpk P). The linear normalization with respect to acoustic pressure magnitude (and steady state (SS) FS) led to a DFW unaffected by strong internal resonances. For the C2H4/N2-air system, from 8 to 20 Hz, DFW is constant at 8.52 plus or minus 0.20 (% weakening)/Pa. This reflects a quasi-steady flame response to an acoustically induced dU(sub air)/dP. Also, it is surprisingly independent of C2H4/N2 mole fraction due to normalization by SS FS. From 20 to approximately 150 Hz, the C2H4/N2 air-flames weakened progressively less, with an inflection at approximately 70 Hz, and became asymptotically insensitive (DFW approximately 0) at approximately 300 Hz, which continued to 1600 Hz. The DFW of CH4-air flames followed a similar pattern, but showed much greater weakening than C2H4/N2-air flames; i.e., the quasi-steady DFW (8 to approximately 15 Hz) was 44.3 %/Pa, or approximately 5x larger, even though the 0 Hz (SS) FS was only 3.0 x smaller. The quasi-steady DFW's of C3H8-air and C2H6-air were intermediate at 34.8 and 20.9 %Pa, respectively. The DFW profiles of all four fuels, at various frequencies, correlated well but non-linearly with respective SS FS's. Notably, the DFW profile for C3H8 air fell more rapidly in the range greater than 15 to 60 Hz, compared with the 1- and 2-carbon fuels. This may indicate a shift in chemical kinetics, and/or O2 transport to a flame that moved closer to the fuel-side. In conclusion, Dynamic Flame Weakening limits appear significant and unique for each fuel, and correlate closely, but non-linearly, with Steady-State Flame Strengths at any given frequency. For reasons unknown, the dynamic flames didn't weaken more at intermediate frequencies (e.g., at 20-50 Hz) than they did at low frequencies (less than 15 Hz), where quasi-steady weakening appears to dominate. Quasi-steady flame weakening ostensibly represents a transient input strain rate maximum that just exceeds the steady-state strain-rate-limited extinction limit for a few cycles. Clearly, further detailed mechanistic understanding is needed in the fall-off region.

Pellett, G.↗
NASA NTRS1994

DNA-DSB in CHO-K1 cells induced by heavy-ions: Break rejoining and residual damage (GSI)

DNA double strand breaks (DSB's) are the critical lesions involved in cellular effects of ionizing radiation. Therefore, the evaluation of DSB induction in mammalian cells after heavy ion irradiation is an essential task for the assessment of high-LET radiation risk in space. Of particular interest has been the question of how the biological efficiency for the cellular inactivation endpoint relates to the initial lesions (DSBs) at varying LETs. For cell killing, an increased Relative Biological Efficiency (RBE) has been determined for highLET radiation around 100-200 keV/mu m. At higher LET, the RBE's decrease again to values below one for the very heavy particles. At GSI, DSB-induction was measured in CHO-K1 cells following irradiation with accelerated particles covering a wide LET range. The electrophoretic elution of fragmented DNA out of agarose plugs in a constant electrical field was applied for the detection of DSB's. The fraction of DNA retained was determined considering the relative intensities of ethidium bromide fluorescence in the well and in the gel lane. Dose-effect curves were established, from which the RBE for DSB induction was calculated at a fraction of 0.7 of DNA retained In summary, these rejoining studies are in line with an enhanced severity of the DNA DSB's at higher LET's, resulting in a decreased repairability of the induced lesions. However, no information concerning the fidelity of strand breaks rejoining is provided in these studies. To assess correct rejoining of DNA fragments an experimental system involving individual DNA hybridization bands has been set up. In preliminary experiments Sal I generated DNA fragments of 0.9 Mbp were irradiated with xrays and incubated for repair However, restitution of the original signals was not observed, probably due to the high radiation dose necessary for breakage of a fragment of this size. A banding pattern with NotI hybridization signals in a higher MW range (3Mbp) has been obtained by varying the electrophoretic conditions and correct rejoining studies will be further developed in this system.

Taucher-Scholz, G.↗
NASA NTRS2004

Metadata Authoring with Versatility and Extensibility

NASA's Global Change Master Directory (GCMD) assists the scientific community in the discovery of and linkage to Earth science data sets and related services. The GCMD holds over 13,800 data set descriptions in Directory Interchange Format (DIF) and 700 data service descriptions in Service Entry Resource Format (SERF), encompassing the disciplines of geology, hydrology, oceanography, meteorology, and ecology. Data descriptions also contain geographic coverage information and direct links to the data, thus allowing researchers to discover data pertaining to a geographic location of interest, then quickly acquire those data. The GCMD strives to be the preferred data locator for world-wide directory-level metadata. In this vein, scientists and data providers must have access to intuitive and efficient metadata authoring tools. Existing GCMD tools are attracting widespread usage; however, a need for tools that are portable, customizable and versatile still exists. With tool usage directly influencing metadata population, it has become apparent that new tools are needed to fill these voids. As a result, the GCMD has released a new authoring tool allowing for both web-based and stand-alone authoring of descriptions. Furthermore, this tool incorporates the ability to plug-and-play the metadata format of choice, offering users options of DIF, SERF, FGDC, ISO or any other defined standard. Allowing data holders to work with their preferred format, as well as an option of a stand-alone application or web-based environment, docBUlLDER will assist the scientific community in efficiently creating quality data and services metadata.

Pollack, Janine↗
NASA NTRS2016

Demonstration of a Packaged Capacitive Pressure Sensor System Suitable for Jet Turbofan Engine Health Monitoring

In this paper, the development and characterization of a packaged pressure sensor system suitable for jet engine health monitoring is demonstrated. The sensing system operates from 97 to 117 MHz over a pressure range from 0 to 350 psi and a temperature range from 25 to 500 deg. The sensing system consists of a Clapp-type oscillator that is fabricated on an alumina substrate and is comprised of a Cree SiC MESFET, MIM capacitors, a wire-wound inductor, chip resistors and a SiCN capacitive pressure sensor. The pressure sensor is located in the LC tank circuit of the oscillator so that a change in pressure causes a change in capacitance, thus changing the resonant frequency of the sensing system. The chip resistors, wire-wound inductors and MIM capacitors have all been characterized at temperature and operational frequency, and perform with less than 5% variance in electrical performance. The measured capacitive pressure sensing system agrees very well with simulated results. The packaged pressure sensing system is specifically designed to measure the pressure on a jet turbofan engine. The packaged system can be installed by way of borescope plug adaptor fitted to a borescope port exposed to the gas path of a turbofan engine.

Harsh environment↗
NASA NTRS2014

AWIPS II Application Development, a SPoRT Perspective

The National Weather Service (NWS) is deploying its next‐generation decision support system, called AWIPS II (Advanced Weather Interactive Processing System II). NASA's Short‐term Prediction Research and Transition (SPoRT) Center has developed several software 'plug‐ins' to extend the capabilities of AWIPS II. SPoRT aims to continue its mission of improving short‐term forecasts by providing NASA and NOAA products on the decision support system used at NWS weather forecast offices (WFOs). These products are not included in the standard Satellite Broadcast Network feed provided to WFOs. SPoRT has had success in providing support to WFOs as they have transitioned to AWIPS II. Specific examples of transitioning SPoRT plug‐ins to WFOs with newly deployed AWIPS II systems will be presented. Proving Ground activities (GOES‐R and JPSS) will dominate SPoRT's future AWIPS II activities, including tool development as well as enhancements to existing products. In early 2012 SPoRT initiated the Experimental Product Development Team, a group of AWIPS II developers from several institutions supporting NWS forecasters with innovative products. The results of the team's spring and fall 2013 meeting will be presented. Since AWIPS II developers now include employees at WFOs, as well as many other institutions related to weather forecasting, the NWS has dealt with a multitude of software governance issues related to the difficulties of multiple remotely collaborating software developers. This presentation will provide additional examples of Research‐to‐Operations plugins, as well as an update on how governance issues are being handled in the AWIPS II developer community.

Burks, Jason E.↗
NASA NTRS1991

Divergence thrust loss calculations for convergent-divergent nozzles: Extensions to the classical case

The analytical derivations of the non-axial thrust divergence losses for convergent-divergent nozzles are described as well as how these calculations are embodied in the Navy/NASA engine computer program. The convergent-divergent geometries considered are simple classic axisymmetric nozzles, two dimensional rectangular nozzles, and axisymmetric and two dimensional plug nozzles. A simple, traditional, inviscid mathematical approach is used to deduce the influence of the ineffectual non-axial thrust as a function of the nozzle exit divergence angle.

Berton, Jeffrey J.↗
NASA NTRS1993

Prediction of prompt NO(x) in hydrocarbon air flames

The gas turbine industry is directing particular attention to very low NOx combustors, whether for aircraft or land based CCGT systems. These low NOx combustors frequently use liquid fuels or natural gas burning under very lean premixed conditions with air or under rich-lean conditions, although only the first case is studied here. In land based systems, diluted steam or nitrogen are sometimes injected into the combustion chambers to reduce flame temperature. The NOx emissions from such systems are the product of three chemical mechanisms which are interrelated: the hydrocarbon prompt NO, the thermal NO (extended Zeldovich mechanism), and the nitrous oxide route to NO. Formation of NO2 from NO also occurs, as well as emission of carbon monoxide and unburnt hydrocarbons. When the fuel-oxidant proportion decreases towards leaner conditions, flame temperatures are lowered, resulting in the total NO being reduced and the thermal-NO contribution greatly diminished to the benefit of the remaining two mechanisms of NO formation. While knowledge of the elementary reactions and their chemical kinetics concerning methane and simple hydrocarbons combustion has existed for a number of years, its use for computer modeling is limited to simple flow dynamics configurations. Nevertheless, understanding of such combusting flows under a wide range of experimental conditions allows for analogies or speculations with more complex actual systems. Such understanding can be achieved by means of one dimensional laminar premixed flame modeling, with a full chemical mechanism which incorporates the three routes of NO formation. Complementary to this understanding is the modeling of the actual combustion system using a full description of the fluid dynamics coupled with a reduced chemical scheme, which is then compared against the first model. The objective of this investigation is to evaluate the relative importance of the three mechanisms of NO formation in lean premixed methane-air combustion with increasing pressure using the one dimensional plug flow package PREMIX, and to test the validity of a three dimensional model with a global chemical mechanism against the one dimensional model in the atmospheric pressure case. Methane is chosen because it is the only mechanism which is reasonably well known and is a good guide to the behavior of other hydrocarbons. The mixture ratio chosen is richer than that in lean gas turbines, but the combustion of this mixture with the low preheat gives realistic gas turbine final flame temperatures. Conditions of NO2 formation are also analyzed in the one dimensional model and results extrapolated to the case of gas turbines.

Valerie Dupont↗
NASA NTRS1980

Time dependent difference theory for sound propagation in axisymmetric ducts with plug flow

The time dependent governing acoustic-difference equations and boundary conditions are developed and solved for sound propagation in an axisymmetric (cylindrical) hard wall duct with a plug mean flow and spinning acoustic modes. The analysis begins with a harmonic sound source radiating into a quiescent duct. This explicit iteration method then calculates stepwise in real time to obtain the transient as well as the 'steady' state solutions of the acoustic field. The time dependent finite difference analysis has two advantages over the steady state finite difference and finite element techniques: (1)the elimination of large matrix storage requirements, and (2)shorter solution times under most conditions.

Baumeister, K. J.↗
NASA NTRS1980

Time dependent difference theory for sound propagation in axisymmetric ducts with plug flow

The time dependent governing/acoustic-difference equations and boundary conditions are developed and solved for sound propagation in an axisymmetric (cylindrical) hard wall duct with a plug mean flow and spinning acoustic modes. The analysis begins with a harmonic sound source radiating into a quiescent duct. This explicit iteration method then calculates stepwise in real time to obtain the transient as well as the 'steady' state solutions of the acoustic field. The time dependent finite difference analysis has two advantages over the steady state finite difference and finite element techniques: (1) the elimination of large matrix storage requirements, and (2) shorter solution times under most conditions.

Baumeister, K. J.↗
NASA NTRS2011

STRS SpaceWire FPGA Module

An FPGA module leverages the previous work from Goddard Space Flight Center (GSFC) relating to NASA s Space Telecommunications Radio System (STRS) project. The STRS SpaceWire FPGA Module is written in the Verilog Register Transfer Level (RTL) language, and it encapsulates an unmodified GSFC core (which is written in VHDL). The module has the necessary inputs/outputs (I/Os) and parameters to integrate seamlessly with the SPARC I/O FPGA Interface module (also developed for the STRS operating environment, OE). Software running on the SPARC processor can access the configuration and status registers within the SpaceWire module. This allows software to control and monitor the SpaceWire functions, but it is also used to give software direct access to what is transmitted and received through the link. SpaceWire data characters can be sent/received through the software interface, as well as through the dedicated interface on the GSFC core. Similarly, SpaceWire time codes can be sent/received through the software interface or through a dedicated interface on the core. This innovation is designed for plug-and-play integration in the STRS OE. The SpaceWire module simplifies the interfaces to the GSFC core, and synchronizes all I/O to a single clock. An interrupt output (with optional masking) identifies time-sensitive events within the module. Test modes were added to allow internal loopback of the SpaceWire link and internal loopback of the client-side data interface.

Lux, James P.↗
NASA NTRSDate not supplied

Study of a Plug Nozzle for Supersonic Aircraft Concepts

Nozzles with external plugs are candidates for the exhaust system of supersonic aircraft concepts in an ongoing research effort at NASA. As part of that effort a fundamental study is being conducted at NASA Glenn Research Center (GRC). Experimental and limited numerical simulation results from this study, for a plug nozzle with a convergent nozzle (cowl),have been reported in recent publications[1,2]. Conventional method-of-characteristics (MoC) design for the plug requires that the flow is choked at the nozzle exit. This is why a sharp convergence of the cowl is invoked in such a design in most previous studies [3-5], leading to the earlier choice of the convergent cowl. Such a nozzle, however, is not suitable for supersonic flights because of large boat tail drag suffered by its outer surface. The boat tail drag consideration dictates that the cowl be more or less cylindrical in shape. While various other shapes have been studied in the past [6], essentially a cylindrical shape is adopted in the larger research program at NASA. The cylindrical shape has certain advantages. With the plug crown located somewhat inside the nozzle, the flow experiences convergence and divergence as with a C-D nozzle. By translating the plug, different throat-to-exit area ratio could be achieved to obtain fully expanded condition at different values of NPR. This offers a relatively easy active control strategy for the nozzle geometry throughout the flight regime from landing and takeoff (LTO) to cruise conditions. A plug nozzle model mimicking the geometry being considered in the NASA program, i.e., having a cylindrical exit, have been fabricated recently for continued fundamental studies. Some of the plugs used with the earlier convergent cowl [1] are readily adopted to this model. In addition, a few other plugs have been obtained following a numerical optimization study [7]. The various configurations are to be investigated for thrust, noise and flow fields, with complementing numerical simulations. Thrust data for a variety of configurations as well as limited noise data are in hand. Further explorations including schlieren visualization are to be conducted in the near future. The proposed paper is to summarize all these results. In the following, a few key results obtained so far are described.

jets↗
NASA NTRS2016

A Structured, Model-Based Systems Engineering Methodology for Operations System Design

Two widely accepted techniques for lowering the cost and risk of developing systems are (1) the use of a defined systems engineering (SE) process or methodology and (2) the reuse of existing (previously built) system components. The first technique is represented, for example, in materials published by NASA (e.g., NASA Systems Engineering Handbook) or by professional societies such as INCOSE (International Council on Systems Engineering). Well-formed SE techniques provide value by establishing the proper scope of the system (e.g., requirements), and by identifying and resolving problems relatively early in project lifecycles, when fixes are less expensive. The second technique (reuse) is applied most commonly to hardware and software; it seeks to avoid replicating design and implementation costs while also reducing risk by placing proven capabilities into operational use. In this paper, we outline a methodology combining these two techniques and extending reuse beyond hardware and software to foundational aspects of a Mission Operation System’s (MOS) design. We describe the system design artifacts that result (e.g., requirements, design documentation), as well as the reusable patterns and elements of the design, and their interrelationships. This approach is enabled by model-based systems engineering (MBSE) techniques and tools and is currently available in SysML form as a plug-in to MagicDraw. Additionally, usage of a rigorous MBSE approach allows for training materials and tutorials to be packaged within the overall model itself. The results of such an approach include decreased cost and risk during the design phase, improved ability of the MOS development team to investigate trade spaces and identify impacts to important flight-ground trade studies. Such results extend into decreased costs and risk in later phases due to improved design, decreased need for late fixes or development of "glue-ware" or scripts to fill unanticipated gaps in functionality, and improved ability to identify and plan testing and other validation activities. Finally, lower operational costs can be expected, both due to improved quality of the MOS, increased ease of maintaining updated knowledge of system configuration, and the fact that training and procedural materials are also updated at the same time as accepted system changes.

Bindschadler, Duane L.↗
NASA NTRS1984

Proteus - An experimenter's view

The scientific experiments package to be carried by the Proteus system takes the form of an Instrument Load carried into orbit by a Space Shuttle, and there mated to a Proteus spacecraft with the Shuttle's Remote Manipulator System. The Proteus system extends to ground support equipment, development tools, and communications, as well as the orbiting satellites. It is expected that Proteus will be able to triple the number of Explorer missions while staying within the current budgetary allocation for such missions. The Proteus spacecraft encompasses a system interface assembly plug, a data handling module, remote interface units, and a power distribution module.

Hibbard, W. D.↗
NASA NTRS1990

The structure and evolution of radiatively cooling jets

The two-dimensional simulations presently used to characterize the structure and evolution of radiatively cooling supersonic jets reveal that cooling jet morphologies resemble those of adiabatic outflows, but with the fundamental difference that a dense, cold shell will condense out of the shocked gas at the head of the jet when the cooling distance behind either of the two principal shocks is smaller than the jet radius. For very high cooling rates, the material that accumulates at the head of the jet forms an extended plug of cold gas resembling the 'nose cone' observed in numerical simulations of strongly magnetized adiabatic jets. An investigation is made of the dependence of jet properties on the density ratio between the beam and the ambient medium, as well as on the strength of radiative cooling.

Blondin, John M.↗
NASA NTRS2012

Model-Based Verification and Validation of Spacecraft Avionics

Our simulation was able to mimic the results of 30 tests on the actual hardware. This shows that simulations have the potential to enable early design validation - well before actual hardware exists. Although simulations focused around data processing procedures at subsystem and device level, they can also be applied to system level analysis to simulate mission scenarios and consumable tracking (e.g. power, propellant, etc.). Simulation engine plug-in developments are continually improving the product, but handling time for time-sensitive operations (like those of the remote engineering unit and bus controller) can be cumbersome.

Systems Modeling Language (SysML↗
NASA NTRS2011

Simulation and Optimization of Vacuum Swing Adsorption Units for Spacesuit Carbon Dioxide and Humidity Control

Controlling carbon dioxide (CO2) and humidity levels in a spacesuit is critical to ensuring both the safety and comfort of an astronaut during extra-vehicular activity (EVA). Traditionally, this has been accomplished utilizing either non-regenerative lithium hydroxide (LiOH) or regenerative but heavy metal oxide (MetOx) canisters which pose a significant weight burden. Although such technology enables air revitalization, the volume requirements to store the waste canisters as well as the mass to transport multiple units become prohibitive as mission durations increase. Consequently, motivation exists toward developing a fully regenerative technology for spacesuit environmental control. The application of solid amine materials with vacuum swing adsorption technology has shown the capacity to control CO2 while concomitantly managing humidity levels through a fully regenerative cycle eliminating constraints imposed with the traditional technologies. Prototype air revitalization units employing this technology have been fabricated in both a rectangular and cylindrical geometry. Experimental results for these test articles have been collected and are described herein. In order to accelerate the developmental efforts, an axially-dispersed plug flow model with an accompanying energy balance has been established and correlated with the experimental data. The experimental and simulation results display good agreement for a variety of flow rates (110-170 ALM), replicated metabolic challenges (100-590 Watts), and atmosphere pressures under consideration for the spacesuit (248 and 760 mm Hg). The testing and model results lend insight into the operational capabilities of these devices as well as the influence the geometry of the device has on performance. In addition, variable metabolic profiles were imposed on the test articles in order to assess the ability of the technology to transition to new metabolic conditions. The advent of the model provides the capacity to apply computer-aided engineering practices to support the ongoing efforts to optimize and mature this technology for future application to space exploration.

Swickrath, Michael J.↗
NASA NTRS1987

Determination of the cross-sectional temperature distribution and boiling limitation of a heat pipe

A computer model is developed and verified which is capable of determining the cross-sectional temperature distribution within a heat pipe with an attached radiator fin; such heat pipes would be plugged into contact heat exchangers designed to carry heat from a space station habitation module to the radiator elements through a centralized fluid loop. The model can furnish information for determining the susceptibility of the monogroove heat pipe to boiling, as well as the location and magnitude of that boiling.

Peterson, G. P.↗
NASA NTRS1994

The Prediction of Nozzle Performance and Heat Transfer in Hydrogen/Oxygen Rocket Engines with Transpiration Cooling, Film Cooling, and High Area Ratios

An advanced engineering computational model has been developed to aid in the analysis of chemical rocket engines. The complete multispecies, chemically reacting and diffusing Navier-Stokes equations are modelled, including the Soret thermal diffusion and Dufour energy transfer terms. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film-cooled nozzle, and a transpiration-cooled plug-and-spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 nozzle and the film-cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. It is demonstrated that thermal diffusion has a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle and was shown to represent a significant fraction of the diffusion fluxes occurring in the transpiration-cooled rocket engine.

Kacynski, Kenneth J.↗